ISS Russian Segment User Manual

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S.P.Korolev Rocket and Space Corporation Energia
ISS Russian Segment
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User Manual
ОАО «РКК «Энергия» им. С.П. Королева», 2004
2011
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Table of Contents
1 Introduction………………………………………………………………………....3
2 Ground rules………………………………………………………………………...4
3 General data on the ISS RS resources at different stages of its deployment……5
3.1 ISS RS consisting of SM and DC1……………………………………………….5
3.2 ISS RS consisting of SM, DC1 and MRM2……………………………………...8
3.3 ISS RS consisting of SM, DC1, MRM2 and MRM1…………………………..11
3.4 ISS RS consisting of SM, MRM2, MRM1 and MLM…………………………14
3.5 Characteristics of communication links and downlinking……………………17
3.6 Trajectory and navigation support for space experiments…………………...19
3.7 The list of photographic and video equipment onboard ISS RS……………..22
3.8 Equipment for still photography Specifications……………………………….23
3.9 Video equipment………………………………………………………………....30
4 Capabilities supporting scientific equipment operation on ISS RS modules….37
4.1 Service module (СМ)……………………………………………………………37
4.2 Docking compartment No.1 (DC1)……………………………………………..54
4.3 Mini Research Module 2 (MRM2)……………………………………………...58
4.4 Mini Research Module 1 (MRM1)……………………………………………...66
4.5 Multipurpose laboratory module (MLM)……………………………………...70
5 Cargo certification………………………………………………………………...89
5.1 Documents used in the course of certification…………………………………90
5.2 Scientific equipment tests……………………………………………………….92
6 Specifications for experimental equipment delivered to ISS………………….102
6.1 Specifications for equipment transported in Russian spacecraft Progress and
Soyuz………………………………………………………………………………...102
6.2 Specifications for the equipment stored and operated onboard ISS RS……120
6.3 Safety requirements……………………………………………………………150
6.4 A typical crew day plan during implementation of the standard mode……159
6.5 Acoustic environment in the ISS RS modules………………………………..160
6.6 Cargo integration………………………………………………………………166
Attachment A……………………………………………………………………….186
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1 Introduction
This Manual provides basic information that designers of experiments and payloads
that are proposed to be conducted and installed in the Russian Segment of the
International Space Station (ISS RS) need to know in order to correctly state
specifications when drawing up statements of work for space experiments and scientific
equipment, as well as to conduct a preliminary feasibility analysis of the experiments
proposed for ISS RS.
It provides general information about ISS RS resources at different stages of its
deployment, as well as about resources of its individual elements. Special attention is
given to the issues related to payload delivery and in-orbit operation and returning
experimental results to earth.
All the payloads installed on ISS RS (Scientific Equipment (SE)) are included into the
Mission Payload Facility (MPF). The Manual describes the ISS RS MPF architecture
and provides a brief description of its components.
It provides a description of workstations for accommodating scientific equipment,
SE, the principles of SE control, and methods of downlinking data from the SE.
It describes the SE operational environments within the ISS RS modules and within
the transportation vehicles Soyuz and Progress.
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mission payloads that support the conduct of space experiments, resources provided for
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2 Ground rules
To meet the necessary requirements for SE development, testing, delivery,
accommodation and functioning, a Mission Payload Facility is established onboard ISS
RS. Included in the MPF are:
- scientific equipment;
- mechanical adapters;
- science support equipment;
- cables.
In order to achieve the most efficient use of resources provided by the space station
for conducting research throughout its life (during which time some research objectives
may become irrelevant, the research methodology may change, and scientific equipment
may fail or become obsolete), the MPF is based on the principles of unification and
integratedness.
The unification principle calls for the use of replaceable payload technology which
implies that virtually any kind of scientific equipment can be delivered and put into
support installation and adaptation of delivered SE (built to meet the relevant
requirements) from the standpoint of mechanical, electrical, data and other interfaces.
SE control, data support, temperature control and vacuumization onboard ISS RS are
provided by the onboard support systems that are not included into MPF.
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operation immediately in orbit. In order to implement such technology, the MPF will
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3 General data on the ISS RS resources at different stages of its deployment
3.1 ISS RS consisting of SM and DC1
Resources allocated to the MPF include:
- the working volume allocated (at SM and DC1): up to 1 m3;
- power supply: up to 0.3 kW (daily average) and up to 1 kW for 2 days with continuous
additional normal power supply to ISS RS from the US segment in accordance with
operational documentation;
- heat removal capacity: up to 0.3 kW (daily average) (the balance is achieved taking
into account the heat loss through the SM hull);
- the number of workstations on the outer surface: 4 pcs. on the SM (8 pcs., when
deliverable Multi-Purpose Workstations (MPWs) are taken into account)
- payload data downlink capacity (maximum, via TM IU unit): up to 2 MB per a
session;
- ISS attitude control mode (nominal): Orbital Coordinate System (OCS);
- ISS OCS attitude control accuracy: ± 10 angular minutes;
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- ISS angular rate stabilization accuracy: 0.005 deg/s;
- the number of telemetered parameters: up to 212;
- the list of electrical interfaces available for MPF:
-RS-232;
-RS-422;
-RS-485;
-Ethernet;
-USB;
- HF with wave impedance coefficients: 75 Ohm and 50 Ohm;
- discrete control commands;
- the number of vaccumization interfaces for MPF: 1 pc.
The MPF cargo traffic to ISS RS is planned based on the following assumptions:
- delivery of cargo to ISS RS onboard Progress M (M1) logistics spacecraft on average
up to 600 kg per year;
- return from ISS RS of up to 110 kg of payloads per year.
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The external appearance of ISS RS in the above configuration is shown in Fig.3.1
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SM
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FGB
Fig. 3.1 – ISS RS configuration consisting of SM and DC1
Progress-M
DC1
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Progress-M
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Soyuz TMA
3.2 ISS RS consisting of SM, DC1 and MRM2
Resources allocated to the MPF include:
- allocated working volume of up to 1 m3 on the SM and up to 0.2 m3 on MRM2;
- power supply:
- on SM and DC1 up to 0.3 kW (daily average) and up to 1 kW for 2 days with
continuous additional normal power supply to ISS RS from the US segment in
accordance with operational documentation;
- on MRM2 up to 0.1 kW (daily average);
- heat rejection capacity: up to 0.3 kW (daily average) (the balance is achieved taking
into account the heat loss through the SM hull);
- the number of workstations on the outer surface:
-4 pcs. on the SM (8 pcs., when deliverable Multi-Purpose Workstations (MPWs) are
taken into account)
-5 pcs. on MRM2 (7 pcs., when deliverable Multi-Purpose Workstations (MPWs) are
taken into account);
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- payload data downlink capacity (maximum, via TM IU unit): up to 2 MB per a
session;
- ISS attitude control mode (nominal): Orbital Coordinate System (OCS);
- ISS OCS attitude control accuracy : ± 10 angular minutes;
- ISS angular rate stabilization accuracy: 0.005 deg/s;
- the number of telemetered parameters: up to 212 (SM) and up to 50 (MRM2);
- the list of electrical interfaces available for MPF:
-RS-232;
-RS-422 (only on SM);
-RS-485 (only on SM);
-Ethernet;
-USB;
- HF with wave impedance of 75 Ohm and 50 Ohm (only on SM);
- discrete control commands;
- the number of vaccumization interfaces for MPF: 1 pc. (on MRM2).
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The MPF cargo traffic to ISS RS is planned based on the following assumptions:
- delivery of cargo to ISS RS onboard Progress M (M1) logistics spacecraft: on average
up to 600 kg per year;
- return from ISS RS of up to 110 kg of payloads per year.
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The external appearance of ISS RS in the above configuration is shown in Fig.3.2
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Progress M
DC1
SM
Progress M
Fig. 3.2 – ISS RS configuration consisting of SM, DC1 and MRM2
MRM2
Soyuz TMA
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3.3 ISS RS consisting of SM, DC1, MRM2 and MRM1
Resources allocated to the MPF include:
- allocated working volume of : up to 1 m3 on SM, up to 0.2 m3 on MRM2 and up to 3
m3 on MRM1;
- power supply:
- on SM and DC1 up to 0.3 kW (daily average) and up to 1 kW for 2 days with
continuous additional normal power supply to ISS RS from the US segment in
accordance with operational documentation;
- on MRM2 up to 0.1 kW (daily average);
- on MRM1 up to 0.1 kW (daily average);
- heat rejection capacity:
- on SM, DC1 and MRM2 up to 0.3 kW (daily average) (the balance is achieved taking
into account the heat loss through the SM hull);
- on MRM1 up to 0.1 kW (daily average);
- the number of workstations on the outer surface:
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-4 pcs. on the SM (8 pcs., when deliverable Multi-Purpose Workstations (MPWs) are
taken into account);
-5 pcs. on MRM2 (7 pcs., when deliverable Multi-Purpose Workstations (MPWs) are
taken into account);
- payload data downlink capacity (maximum, via TM IU unit): up to 2 MB per a
session;
- ISS attitude control mode (nominal): Orbital Coordinate System (OCS);
- ISS OCS attitude control accuracy : ± 10 angular minutes;
- ISS angular rate stabilization accuracy: 0.005 deg/s;
- the number of telemetered parameters: up to 212 (SM), up to 50 (MRM2) and up to 36
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(MRM1);
- the list of electrical interfaces available for MPF:
-RS-232;
-RS-422 (only on SM);
-RS-485 (only on SM);
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-Ethernet;
-USB;
- HF with wave impedance of 75 Ohm and 50 Ohm (only on SM);
- discrete control commands (on SM and MRM2);
- the number of vaccumization interfaces for MPF: 1 pc on MRM2 and 1 pc on MRM1.
The MPF cargo traffic to ISS RS is planned based on the following assumptions:
- delivery of cargo to ISS RS onboard Progress M (M1) logistics spacecraft: on average
up to 600 kg per year;
- return from ISS RS of up to 110 kg of payloads per year.
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The external appearance of ISS RS in the above configuration is shown in the figure
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DC1
Soyuz TMA
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SM
Progress M
Progress M
MRM2
Fig. 3.3 – ISS RS configuration consisting of SM, DC1, MRM2 and MRM1
Soyuz TMA
FGB
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MRM1
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3.4 ISS RS consisting of SM, MRM2, MRM1 and MLM
Resources allocated to the MPF include:
- allocated working volume of : up to 1 m3 on SM, up to 0.2 m3 on MRM2, up to 3 m3
on MRM1, and up to 8 m3 on MLM;
- power supply:
- on SM and DC1 up to 0.3 kW (daily average) and up to 1 kW for 2 days with
continuous additional normal power supply to ISS RS from the US segment in
accordance with operational documentation;
- on MRM2 up to 0.1 kW (daily average);
- on MRM1 up to 0.1 kW (daily average);
- on MLM up to 1.0 kW (daily average) inside the pressurized cabin and up to 1.5 kW
(daily average) outside the pressurized cabin;
- heat rejection capacity:
- on SM, DC1 and MRM2 up to 0.3 kW (daily average);
- on MRM1 up to 0.1 kW (daily average);
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- on MLM up to 1.0 kW (daily average) after deployment of an additional radiative
heat exchanger;
- the number of workstations on the outer surface:
-4 pcs. on the SM (8 pcs., when deliverable Multi-Purpose Workstations (MPWs) are
taken into account);
-5 pcs. on MRM2 (7 pcs., when deliverable Multi-Purpose Workstations (MPWs) are
taken into account);
-9 pcs. on the MLM (13 pcs., when deliverable Multi-Purpose Workstations (MPWs)
are taken into account);
- payload data downlink capacity (maximum, via TM IU unit): up to 2 MB per a
session;
- ISS attitude control mode (nominal): Orbital Coordinate System (OCS);
- ISS OCS attitude control accuracy : ± 10 angular minutes;
- ISS angular rate stabilization accuracy: 0.005 deg/s;
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- the number of telemetered parameters: up to 212 (SM), up to 50 (MRM2), up to 36
(MRM1), up to 323 (MLM);
- the list of electrical interfaces available for MPF:
-RS-232;
-RS-422 (on SM and MLM);
-RS-485 (on SM and MLM);
-Ethernet;
-USB;
-MIL STD 1553B (only on MLM);
- HF with wave impedances of 75 Ohm and 50 Ohm (on SM and MLM);
- discrete control commands (on SM, MRM2 and MLM);
- the number of vaccumization interfaces for MPF: 1 pc on MRM2, 1 pc on MRM1, 2
pcs on MLM;
- the number of thermostatting interfaces for MPF: 2 pcs. on MLM.
- delivery of cargo to ISS RS onboard Progress M (M1) logistics spacecraft: on average
up to 600 kg per year;
- return from ISS RS of up to 110 kg of payloads per year.
The external appearance of ISS RS in the above configuration is shown in the Figure
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The MPF cargo traffic to ISS RS is planned based on the following assumptions:
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SM
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MLM
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FGB
MRM2
Progress M
Soyuz TMA
MRM1
Fig. 3.4 – ISS RS configuration consisting of SM, MRM2, MRM1 and MLM
Progress M
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3.5 Characteristics of communication links and downlinking
The preliminary design calls for establishing on ISS RS intercomputer digital data
exchange links with MCC-M via RF links of the standard SM systems – Telephone and
Telegraph Communications System (TTCS), RF communications and control system
(RF C&CS) Regul-OS, TV System (TVS) and onboard telemetry system OI TS 2-12
with data throughput of:
- 1.2 kbps via the Telephone & Telegraph System Interface Unit (TTS IU) and the
TTCS VHF link (up- and down-linking);
- 2.4/4.8 kbps via interface units IU or TM IU and RF-link phone-2 of RF C&CS Regul
OS (up- and down-linking);
-16, 32, 64 kbps via interface units IU or TM IU and RF-link phone-3 of RF C&CS
Regul OS (up- and down-linking);
- 51.2/102.4 kbps via TM IU and RF-link of the OI TS 2-12 system (subsystem B) in
the data array transmission mode (downlinking);
Regul-OS when using Zveno-B hardware (up- and down-linking);
Out of the abovementioned communications links, currently implemented onboard
SM and available for digital data exchange are:
- RF C&CS Regul OS RF-link phone-2 with 4.8 kps capacity (via IU unit);
bandwidth, and for that reason is impractical for digital data transmission and is
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- the TTCS VHF link with 1.2 kbps capacity (via TTS IU unit);
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- 16, 32, 64 kbps (altogether up to 96 kbps) via user channels ПГ 1,2 of RF C&CS
- RF C&CS Regul OS RF-link phone-3 with 16 kps capacity (via IU unit).
The TTCS VHF link is mostly used for communications with the crew, has a low
assigned a backup status.
Exchange of data at 4.8 kbps over the phone-2 channel of RF C&CS Regul OS via the
flight model of IU makes it possible to downlink up to 100 KB of data during a 5minute communications session.
Exchange of data at 16 kbps over the phone-3 channel of RF C&CS Regul OS in the
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forward (uplink) channel is working normally, but the downlink, for the time being,
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provides no more than 150 KB of data over a 5-minute communications session. If
larger files need to be downlinked, a series of several consecutive communication
sessions can be used. Besides, as the next phase in its upgrade, there are plans to switch
to (after proper testing is completed) 32 and 64 kbps transmission modes for the phone3 channel of Regul-OS.
The maximum amount of payload data that can be downlinked via TM IU over
telemetry channel per one communications session is 2 MB, with other users not being
able to use this communications channel while the payload data are being downlinked.
In early 2011, a Radio Data Transmission System (RDTS) is to be deployed on the
SM of the ISS RS, which is to provide additional data downlink capacity:
- RF-link with up to 100 Mbps throughput (downlink);
- receiving data over HotLink from the payload installed on the outer surface of the SM
for its direct transmission.
The RDTS system allows downlinking no less than 3.7 GB of data during a 5-minute
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communications session.
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3.6 Trajectory and navigation support for space experiments
ISS, together with the ISS RS, is continuously (normally) operated in the orbital
coordinate system.
The accuracy and stability of maintaining the ISS in the OCS are:
- ISS OCS attitude control accuracy : ± 10 angular minutes;
- ISS angular rate stabilization accuracy: 0.005 deg/s
Additional requirements for payload accuracy and stabilization during a program of
experimental tests shall be met through the use of additional specialized (or
multipurpose) equipment working in conjunction with the operating mission payload
equipment.
ISS can perform programmed pitch and roll maneuvers within ± 15 degrees (this
constraint is imposed due to the lighting conditions of the thermal control system panels
on the US orbital segment).
Installed at present on one of the on-board crew support computers is the Sigma
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program for displaying the trajectory and navigation flight environment, which makes it
possible to perform the following tasks:
- continuous real-time display on the screen of the space station position, current orbit
track, coverage areas of the nearest ground stations, the start and end times of the next
communications session with Earth, lighting environment in orbit, mission elapsed
time, etc. Used to display the “running view” of the current location under the flight
path are Earth surface maps of various scale: ranging from coarse (mapboard-type
maps) to detailed (1:200000). As all the maps are bitmap images on cylindrical
projection, it is also possible to use any other set of maps or color photos (prepared in
the same format);
- provide real-time simulations of the future (prediction) or the past position of the
space station. Selected as variable simulation parameters can be year, month, day, hour
and second of the flight, as well as the number of the orbit and the number of days into
the mission. When the simulation is run “forward”, the latest available orbital data are
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extrapolated, when the simulation is run “backwards”, the history of orbital and attitude
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data is used. There is a “fast-motion” mode for displaying the space station position. In
addition to this, the space station position can be "frozen” in its orbital track on the
screen at any desired moment in time.
- review all the available electronic maps in the “geographic atlas" mode and determine
geographic coordinates of any area on the Earth surface;
- to automatically determine on the e-maps the names of a region, a country, a water
area (i.e., a sea, a bay, a sound), or an area of the Russian Federation, which is being
overflown by the space station, and output these data in a specified window on the
display screen;
- to promptly calculate (predict) when and how well a specified area on the Earth
surface is going to be visible from the space station within the interval of time covering
the next 16 orbits to a week. Coordinates of a region can be specified by means of an
arbitrary point immediately on the screen itself (on the map) or selected from any preprepared list of objects. The trajectory and navigation support software includes lists on
the following subjects: ecology, geology, etc. Selecting an item on these lists can be
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done by typing the name of the desired item on the keyboard. The visibility can be
calculated for several targets at the same time, with subsequently displaying the data on
what targets are visible during the current orbit. Also calculated are the coverage areas
of the specified ground stations;
- to evaluate reliability (the level of likelihood) of detecting and identifying the known
feature, a group of features or a sector of the terrain in the observed region taking into
account geographic and meteorological observation conditions. The evaluation takes
into consideration the presence of cloud cover (intensity on the scale of 0 to 10) over the
region under observation. The cloud cover intensity is automatically evaluated based on
monthly average values or is set manually based on real-time data. Evaluated
automatically is also the number of observations which need to be planned in order to
identify the desired feature with the specified level of reliability.
- to simulate in real time a segment of the Earth surface and/or of the celestial sphere
which is visible through the specified porthole. To simulate the celestial sphere,
Pulkovo observatory catalog is used – 5050 brithest stars (recalculated for 2003 epoch);
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also continuously calculated are positions of the Sun (and solar reflection from the
Earht surface), the Moon and planets, Other SC, for which orbital parameters are
available. There is a mode for the automatic selection of that porthole, which is the best
suited for observing the specified feature on the Earth surface or the celestial sphere. In
addition to the actual portholes, there is a "virtual porthole”, which simulates the view
forward (or backward) along the flight path.
- to simulate the space station rotation in coordinates bound with the specified window.
This makes it possible, even when there are no available data on the current attitude, to
determine the actual attitude by manually bringing into coincidence the electronic
model of the porthole and the actual image visible through the porthole, using in the
process such visual cues as the horizon line, the Sun, the Moon and bright stars.
to receive voice messages about various events (beginning and end of a
communications session, orbital sunsets and sunrises, a new orbit, etc.).
To assure normal operation of the trajectory and navigation support software, provisions
have been made for daily uplinks of the orbital trajectory data predictions.
to the space stations on removable hard disk drives onboard Soyuz TMA spacecraft.
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New versions of the software with the updates reflecting crew comments are delivered
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3.7. The list of photographic and video equipment onboard ISS RS
Currently available onboard ISS RS is equipment (still photo/video cameras), which
can support, within certain limits, the tasks of Earth remote sensing.
Tables 3.7.1 and 3.7.2 provide the lists of photographic and video equipment onboard
ISS RS, respectively.
Table 3.7.1 The list of photographic equipment onboard ISS RS
Major accessories
Equipment name
Still camera
Nikon D2X
Nikkor autofocus lenses
Still camera
Nikon D3
Nikkor autofocus lenses
Still camera
Nikon D3
Nikkor autofocus lenses
Table 3.7.2 The list of video equipment onboard ISS RS
Equipment name
Major accessories
Video system LIV
Telescopic
lens
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Fujinon A16x9 BRM-28A
HDTV videosystem
Telescopic lens HJ15x8B IAS
HDV videocamera
The camcoder contains a high-resolution
(Sony HVR-Z1 and Sony HVR-Z7)
video camera with 3 CCD matrices and
Video camera Sony DSR PD150P
12x zoom lens matches the 1/3-inch
CCD system
Optical stabilizer - Super SteadyShot™
A more detailed description, characteristics and external appearance of the above
equipment are given in sections 3.8 and 3.9.
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HDV video recorder
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3.8. Equipment for still photography Specifications.
Currently available onboard ISS RS are still cameras Nikon D2X,
Nikon D3 and Nikon D3X (see Table 3.8.1.1), which make it possible to take highresolution photos of the Earth surface and other through ISS portholes, mostly portholes
No.6, No.7 and No.8.
The still cameras come with accessory lenses which are listed below and are intended
for different tasks.
Table 3.8.1.1 Russian photographic equipment onboard SM
No
Name
.
Designation
Qty.
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Still cameras
1
Still camera
Nikon D3X
SM-FOTO-D3X-U01
2
2
Still camera
Nikon D3
SM-FOTO-D3-U01 №01
2
3
Still camera
Nikon D2X
SM-FOTO-D2X-U01
2
Lenses
1
Lens AF DX Nikkor 10.5 mm
SM-FOTO-D1-U02-07
1
2
Lens AF-S Nikkor 14 mm
SM-FOTO-D1X-U02-03
1
3
Lens AF-S Nikkor 17-35 mm
SM-FOTO-D1X-U02-01
1
Lens AF-S Nikkor 17-55 mm
SM-FOTO-D1-U02-06,
4
5
6
7
SM-FOTO-D200-U02 06
Lens AF-S Nikkor 28-70 mm
SM-FOTO-D1X-U02-02
SM-FOTO-F5-1-U02
Lens AF-S Nikkor 80-200 mm
Lens AF VR Zoom-Nikkor
80-400 mm f/4.5-5.6
№ докум.
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2
SM-FOTO-D1X-U02-05
1
SM-FOTO-D3-U02 №2
1
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8
SM-FOTO-D3X-U02-08
1
SM-FOTO-D200-U02-09
1
10 Lens Nikkor 14-24 mm
SM-FOTO-D3X-U02-03
1
11 Telephoto lens AF-S Nikkor 600mm f/4
SM-FOTO-D3X-U02-07
1
12 Telephoto lens AF-S Nikkor 400mm f/2.8 SM-FOTO-D3X-U02-06
1
13 Telephoto lens Nikon TC-14E
SM-FOTO-D3X-TC-14
1
14 Telephoto lens Nikon TC-20E
SM-FOTO-D3X-TC-20
1
15 Telephoto lens Nikon TC-17E
SM-FOTO-D3X-TC-17
1
9
Lens SIGMA AF 300-800 f/5.6
The lens for ultra close-up photography
Micro-Nikkor f = 105 mm
Flash lamps
1
Flash lamp
Nikon SB-800
SM-FOTO-D2X-U03
2
1
2
Battery charger for
camera Nikon D2X
Battery charger for
camera Nikon D200
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Battery chargers
SM-FOTO-D2X-U05
1
SM-FOTO-D200-U05
1
Power supply units
1
Storage battery
Nikon EN-EL4
SM-FOTO-D2X-U06
6
2
Storage battery
Nikon EN-EL4a
SM-FOTO-D2X-U06
2
SM-FOTO-D2X-01
1
Accessories
1
A cable for connecting to a Laptop
computer
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2
Remote control cable
SM-FOTO-D2X-02
3
3
Flash lamp synchronizer СФ-5У
SM-FOTO-D1X-U08
2
SM-FOTO-D1X-U09
2
SM-FOTO-D1X-U10
2
Photographic equipment cleaning kit
SM-FOTO-D1-U12
1
Battery handle for
SM-FOTO-D200-U06-01
4
5
6
7
(for the bracket)
PC CARD ADAPTER
(PCMIA adapter)
camera Nikon D200 MB-D200
Power supply units container
SM-FOTO-D200-U04
2
9
Stereo adapter PENTAX
SM-FOTO-D200-SA
1
11
Adapter
ring
for
stereo
adapter SM-FOTO-D200-SA-R
1
PENTAX
A bracket for stereo imaging through a SM-FOTO-D200-BR
1
porthole
Photographic equipment is self-contained and is not connected to the onboard power
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supply line. Chargers for storage batteries of the Nikon camera are hooked up to
connectors on the LIV video system.
Table 3.8.1.2
Подпись и дата
1
8
10
Подпись и дата
Adapter
Still cameras specifications (Table 3.8.1.2)
Digital still camera
Nikon D2X
CCD matrix dimensions
-
23.7x15.7 mm
Image size
-
4288 x 2848
Light sensitivity
-
Equivalent to 100÷800 ISO speed
Recording medium
-
Compact Flash and Microdrive
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memory cards
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Compression
-
RAW (NEF) and JPEG formats.
LCD display
-
2.5-inch 232000 with adjusted
lighting
Lenses
-
Autofocus lenses Nikkor and SIGMA
AF 300-800 f/5.6
Viewfinder field of view
-
96% of the frame area
Optical exposure
-
Automatic (software controlled with
time of exposure or aperture having
higher priority) and manual
Optical exposure measuring area
-
Matrix centrally weighted spot
measurement
Exposure time range
-
1/8.000÷30s
Bracketing mode
-
2-9 frames with incremental steps of
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1/3, 1/2, 2/3 or 1 EV
Synchronization with the flash
-
5 modes
Power
-
Li-Ion battery EN-EL4a (11.1V)
Battery charger
-
SM-FOTO-D1X-U05
Digital still camera Nikon D3 (Nikon D3X)
CCD matrix dimensions
-
36.0x23.9 mm (35.9х24.0 mm)
Image size
-
4288x2848 (6048x4032)
Light sensitivity
-
Equivalent to 200÷6400 ISO speed
25600 ISO (100÷1600)
Recording medium
-
Compact Flash and Microdrive
memory cards
Compression
-
RAW (NEF) and JPEG formats.
LCD display
-
3-inch, with 920000 pixels, with
adjusted brightness, 100% frame
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coverage
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Lenses
-
Autofocus lenses Nikkor and SIGMA
AF 300-800 f/5.6
Optical exposure
-
Automatic (software controlled with
time of exposure or aperture having
higher priority) and manual
Optical exposure measuring area
-
Matrix centrally weighted spot
measurement
Exposure time range
-
1/8000÷30s
Bracketing mode
-
2-9 frames with incremental steps of
1/3, 2/3 or 1 EV
-
5 modes
Power
-
Li-Ion battery EN-EL4a (11.1V)
Dimensions
-
159.5x157x87.5 mm
Mass
-
1.24 kg (1.22 kg)
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Synchronization with the flash
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Additional characteristics, external appearance and purpose of the above cameras are
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shown in Fig. 3.8.1.1, 3.8.1.2.
Fig. 3.8.1.1 External view of still camera Nikon D3 (Nikon D3X)
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Fig. 3.8.1.2 External view of still camera Nikon D2
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3.9. Video equipment
3.9.1 LIV system specification
LIV system is based on professional equipment operating in analog format
BETACAM
SP.
Fig. 3.9.1.1 shows the video camera included in this system, and Table 3.9.1.1 LIV
system performance data for Earth imaging f rom the altitude of 400 km using the lens
Fujinon A16x9 BRM-28A (the lens is currently onboard ISS RS).
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Fig. 3.9.1.1 Overall view of the LIV system video camera
Table 3.9.1.1 LIV system specification
Parameter
Number of pixels in the CCD matrix
Value
437664 (752x582)
Focal distance (max.), mm
144
Field of view, deg.
3.3x2.38
The imaged area on the ground, km2
23.2х18.4
Resolution per 1 pixel, m
30.9
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At present, the video system does not have the optics, which would meet the quality
requirements for Earth imaging for the purposes of Earth Remote Sensisng.
3.9.2. HDTV video system
Specifications.
Table 3.9.2.1 list HDTV system performance data during Earth imaging from the
altitude of 400 km using the lens HJ15x8B IAS (the lens is currently available onboard
ISS RS).
Table 3.9.2.1 The HDTV system specification
Number of pixels in the CCD matrix
2033600 (1920x1080)
Focal distance (max.), mm
120
240
(without
(with extender)
Field of view, deg.
4.6x2.6
2.3x1.3
The imaged area on the ground, km2
32x18.4
16x9.2
16.6
8.3
Resolution per 1 pixel, m
consultation with the Japanese, it can be used as the ISS RS system equipment. Its
equipment with respect to the LIV system (the resolution of the HDTV camera is 4
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The owner of the HDTV video system is NASDA, however, at present, after
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extender)
further use must be agreed with NASDA, with participation of Rosaviacosmos.
From the technical standpoint, the HDTV video system is the next-generation video
times that of the LIV camera).
In addition to this, HDTV provides more color information about the scene, which is
important during interpretation and analysis of the features on the ground.
The overall view of the HDTV camera is given in Fig. 3.9.2.1.
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It should be noted that the filming results can only be returned to Earth on a recording
medium (HD video cassettes). The existing capacities of the ISS RS RF-links do not
support downlinking of the filming results.
At present, the Russian side has neither the equipment for ground processing of the
information, nor the capabilities to review and copy video materials obtained in the
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course of the HDTV system operation.
Fig. 3.9.2.1 Overall view of the HDTV camera
3.9.3 HDV video system.
The HDV video system is intended for performing the following tasks:
- video recording of the experiments
- video filming inside the SM
- video filming of the Earth
The HDV equipment includes camcorders Sony HVR-Z1 and Sony HVR-Z7 of the
HDV digital format. Using an adapter to the bracket LIV/106/20 the camcorders can be
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installed on the bracket in order to fasten them to the SM handrails. The HDV kit also
includes a wide-angle attachment x0.8 Sony VCL-HG0872, Sony batteries
NP-F970, a lamp Sony HVL-LBP, video cassettes Sony PHDVM-63DM, a set of cables
and adapters.
The list of equipment (Table 3.9.3.1)
Table 3.9.3.1
№
Name
Designation
Qty.
1.
Camcorder Sony HVR-Z1
SM-HDV-U01
1
2.
Camcorder Sony HVR-Z7
SM-HDV-U10
2
3.
Lens Fujinon HAs18x7.6BRM 137mm
SM-HDV-U11
1
4.
Adapter ACM-17
SM-HDV-U12
1
5.
Camera light Sony HVL-LBP
SM-HDV-U14
2
6.
Wide-angle attachment x0.8 Sony VCL-
SM-HDV-U02
1
Camera bag
SM-HDV-U07
1
8.
Camera bag
SM-HDV-U08
1
9.
Battery Sony NP-F970
SM-HDV-U03
6
10. Video cassettes Sony PHDVM-63DM
SM-HDV-U04
14
11. Memory card CF 16Gb
SM-HDV-DCV
8
12. Video-audio cable (1.5m)
SM-HDV-01
1
Взам. инв. №
13. Video-audio cable (1.5m)
SM-HDV-10
1
14. Video cable (5 m)
SM-HDV-02
1
15. Video cable (5 m)
SM-HDV-03
1
16. Adapter RCA-BNC
SM-HDV-04
1
17. Adapter BNC -BNC
SM-HDV-05
1
18. Adapter BNC -BNC
SM-HDV-06
1
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7.
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Подпись и дата
HG0872
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№
Name
19. Adapter to bracket LIV/106/20
Designation
Qty.
From LIV/106/20 to
1
LIV video camera
20. Adapter iLink for MRU
SM-HDV-U15
1
21. Cable iLink 6р-4р (2m)
SM-HDV-11
1
External appearance, characteristics and possible uses of the above video camera are
given in Fig. 3.9.3.
3.9.4 Video camera Sony DSR PD150P
External appearance, characteristics and possible uses of the above video camera are
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given in Fig. 3.9.4.
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Camcoder Sony HVR-Z1 contains a high-resolution video
camera with 3 CCD matrices and HDV video recorder
It can be installed onto a bracket for securing it to the
SM handrails
The camera can be equipped with a wide-angle
attachment x0.8 Sony VCL-HG0872
The camcoder can work in PAL and NTSC systems, as
well as in HDV, DV and DVCAM formats
It is also possible to downlink video and audio signals via
the switch of the LIV video complex
HDV signals are compressed in MPEG2 format used in
ground digital HDTV channels, as well as in Blu-ray disk
recorders
Viewfinder: Electronic (color).Combined 12x (optical)
zoom lens
The camcorder has a system of color filters and balance
of white as well as an amplifier and an electronic shutter
Possible use:
- video recording the progress of the experiments
within experimental setups
- video recording inside SM to provide information
to general public
- video recording of Earth
- video recording from Soyuz during re-docking
operations;
- backup capability for TV broadcasts when
primary LIV system is unavailable or out of order
35
Лист
Fig. 3.9.3 Video camera HDV (Sony HVR-Z1)
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Built-in portable LCD display for viewing the
recordings
The video camera uses Sony’s Memory Stick
The camera image sensor has three 1/3-inch CCD
matrixes with a total of 450 000 pixels
Characteristics of the 12-x zoom lens match 1/3-inch
CCD system
The optical stabilizer is Super SteadyShot™, which
uses independent determination of horizontal and
vertical motion
A prism system in front of the lens effectively
compensates camcordershaking when filming
without a tripod
Black and white LCD viewfinder with 180 000 pixels
provides horizontal resolution of 500 TVL, which
simplifies manual focusing control
Video data are recorded on cassettes. One cassette
holds 40 minutes of recording
Fig. 3.9.4 Video camera Sony DSR PD150P
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36
4 Capabilities supporting scientific equipment operation on ISS RS modules
4.1 Service module (СМ)
The following interfaces with the module are used during integration of the
scientific equipment into the SM:
- mechanical interface;
- electrical interfaces, including power interface, telemetry interface, data
interfaces;
- thermal interface;
- vacuum interface (connection to external environment).
The SM design provides for a number of areas for installing the scientific
equipment in orbit, which have structural support for mechanical attachment and
connections of the scientific equipment to the ISS support systems.
Such areas outside the SM are:
- the УРМ-Н1 workstation, located on the SM outer surface along plane III of the
Подпись и дата
Working Compartment 1 between frames 3, 4. Installed on the SM body on the
transition plate 17КС.300Ю1573-0 as a mechanical device for fastening scientific
equipment if a Passive Base Point (PBP) 27КСМ.152Ю7200-0. Scientific
Equipment is installed by an operator during EVA onto PBP using an adapter,
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which includes an Active Base Point (ABP) 27КСМ.152Ю7100-0;
- workstation УРМ-Н2 formed by a system of handrails 25x25 mm in crosssection on the body of the SM Working Compartment 2 between planes I and II in
such a way that the normal to the line connecting the centers of cross-sections of
the two central handrails is slanted at an angle of 47о from plane I towards plane II.
The scientific equipment is fastened onto the handles with special locks;
- each of the workstations УРМ-Н3 (УРМ-Н4) is formed by four threaded М12
holes with the base of 400×x200 mm, located in the payload fairing supports on
the outer surface of the SM along planes II and IV between frames 8, 8а,
respectively. Scientific equipment is installed using adapters;
Инв.№ подл.
- the Deliverable Multi-Purpose WorkStation (DMPWS) 17КС.600Ю 1501А-0 is
installed on УРМ-Н3. DMPWS consists of a base 17КС.600Ю 1501А-100 with
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PBP and a platform 17КС.600Ю 1501А-200 with three passive parts of scientific
equipment adapters 17КС.600Ю 1150А-0 and a mating device 27КСМ.152Ю
7110-0 for installation onto a PBP. Normals of the three passive adapters installed
on the platform shall be aligned with axes ±YSM and +ZSM.
Installed on the УРМ-Н4 under flight test conditions can be a deliverable
workstation developed with the use of design solutions implemented in the design
of the DMPWS.
The following volumes are reserved for installation and connection of scientific
equipment inside the pressurized cabin of the Service Module:
- on the backside of panel 407 (УРМ4-3 zone) for scientific equipment with
dimensions of up to 175x400x60 mm;
- behind panel 305 (УРМ4-2 zone) for scientific equipment with dimensions up to
220x350x300 mm.
The SM external appearance is shown in Figures 4.1.1 and 4.1.2.
equipment are shown in Figures 4.1.3 and 4.1.4.
SM interior view, locations of portholes, laptops and mission payloads are shown
in Figures 4.1.5, 4.1.6 and 4.1.7.
SM resources used for integration of scientific equipment are listed in Table 4.1.1.
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External multi-purpose workstations on the SM for installing units of scientific
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Fig. 4.1.1 – SM external view
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Fig. 4.1.2 – SM external view
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MPW-N1
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MPW-N2
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MPW-N3
MPW-D
Fig. 4.1.3 – External MPWS for installing scientic equipment units on SM
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Fig. 4.1.4 – External MPWS for installing scientific equipment units on SM
MPW-N4
MPW-D-4
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MPW-N3
MPW-D
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Laptop
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426 mm porthole
228 mm porthole
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Fig. 4.1.5 – SM interior view (starboard)
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Multipurpose biotechnology thermostat
228 mm porthole
Onboard power outlets
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426 mm porthole
Laptop
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Fig. 4.1.6 – SM interior view (port side)
Cryogem-03
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Onboard power outlets
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Fig. 4.1.7 – SM interior view (floor)
Multipurpose biotechnology thermostat
228 mm porthole
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Table 4.1.1 – SM resources for integration of scientific equipment
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Deliverable Multipurpose Workstation
Passive payload adapter
17КС.600Ю 1152А-0
Active payload adapter
17КС.600Ю 1151А-0
Payload adapter
17КС.600Ю1150А-0
Deliverable multipurpose
workstation
17КС.600Ю 1501А-0
Platform with adapters
17КС.600Ю 1501А-210
Passive basepoint
27КСМ.152Ю 7200-0
Retaining plate
(ФИ-21)
17КС.600Ю 1501А-450
Mating device Active Base
Point 27КСМ.152Ю 7110-0
Base
17КС.600Ю 1501А-170
Retaining plate (ФИ-20)
17КС.600Ю 1501А-350
Инв. № дубл.
Подпись и дата
Fig. 4.1.8 – Deliverable multipurpose workstation
Sample fields of view of scientific equipment installed on the SM multipurpose
workstation are given in the following figures:
- Fig. 4.1.9 – field of view of scientific equipment installed on the УРМ-Н1 (axis
of sight aimed at zenith);
- Fig. 4.1.10 – field of view of scientific equipment installed on the УРМ-Н3 (axis
of sight pointed towards zenith);
Инв.№ подл.
Подпись и дата
Взам. инв. №
- Fig. 4.1.11 – field of view of scientific equipment installed on the УРМ-Н3 (axis
of sight pointed towards nadir);
- Fig. 4.1.12 – field of view of scientific equipment installed on the УРМ-Н4 (axis
of sight pointed towards zenith);
- Fig. 4.1.13 – field of view of scientific equipment installed on the УРМ-Н4 (axis
of sight pointed towards nadir);
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Подпись и дата
of sight aimed at zenith).
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
- Fig. 4.1.9 – field of view of scientific equipment installed on the УРМ-Н1 (axis
П40463
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Подпись и дата
of sight pointed at zenith)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
- Fig. 4.1.10 – field of view of scientific equipment installed on the УРМ-Н3 (axis
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Подпись и дата
of sight pointed at zenith)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
- Fig. 4.1.10 – field of view of scientific equipment installed on the УРМ-Н3 (axis
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Подпись и дата
of sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Fig. 4.1.11 – field of view of scientific equipment installed on the УРМ-Н3 (axis
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Подпись и дата
of sight pointed towards zenith)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
- Fig. 4.1.12 – field of view of scientific equipment installed on the УРМ-Н4 (axis
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Подпись и дата
of sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Fig. 4.1.13 – field of view of scientific equipment installed on the УРМ-Н4 (axis
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4.2 Docking compartment No.1 (DC1)
Special features that DC1 provides for conducting scientific experiments are:
- capability to install hardware on the outer surface of MRM2 for observing the
plane of the local horizon;
- availability of a porthole with a capability to sight the hardware in the horizon
plane;
- capability to place hardware inside pressurized cabin for short periods of time for
conducting space experiments and to store it only for the periods between EVAs;
- on-board systems resources for scientific equipment on the external surface are
mostly borrowed from the SM.
On the DC1 external surface, possible areas for deploying payloads are:
- magnetomechanical latches 77КСД.1940-100, located on the cylindrical part and
on the central sphere.
- passive base points 27КСМ.152Ю 7200-0 (2 pcs.) located on the cylindrical part
- brackets 240ГК.А2200А-101 for the latch of the outboard workstation placed (4
pcs. each) on the flange of the exit hatches.
Locations
on
DC1
of
external
multipurpose
workstations
based
on
electromechanical latches are shown in Fig.4.2.1.
Locations on DC1 of payload equipment, windows and electrical outlets are shown
in Fig.4.2.2.
DC1 resources used for integration of scientific equipment are listed in Table 4.2.1.
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
(used for accommodating cargo booms ГСтМ).
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Подпись и дата
Инв. № дубл.
Подпись и дата
Изм. Лист
№ докум.
Подп.
Дата
Fig. 4.2.1 – Locations of external MPWS on DC1
Electrical connector
Micro-meteorid shield 2
Взам. инв. №
Micro-meteorid shield 3
Инв.№ подл.
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Electrical connector
Micro-meteorid shield 1
Инв.№ подл.
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Взам. инв. №
Инв. № дубл.
Onboard power outlet
Подпись и дата
228 mm porthole
Cryogem-03M
Fig. 4.2.2 – Locations of payload equipment, portholes and outlets of the onboard power grid on DC1
228 mm porthole
Подпись и дата
Инв.№ подл.
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Подпись и дата
Инв. № дубл.
Подпись и дата
Table 4.2.1 – DC1 resources for integration of scientific equipment
Взам. инв. №
4.3 Mini Research Module 2 (MRM2)
Characteristics:
- launch mass, kg: 3670±50 kg;
- pressurized volume, m3: 12,5;
- storage volume for cargoes and scientific equipment, m3: 0,2;
- in particular, for scientific equipment, m3: 0,1;
- power for scientific equipment, kW: up to 0.1;
- maximum daily average rejection of heat from the hardware included in the MPF
through the air loop of the thermal control system: up to 0.1;
- launch vehicle: Soyuz-FG
MRM2 external appearance is shown in Fig. 4.3.1
Micrometeorid
shield
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
Advanced cargo boom
(on the passive base point)
Egress hatch (with a porthole)
Deliverable MPWS
(on the passive base point)
Fig. 4.3.1 – MRM2 external appearance.
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Special features that MRM2 provides for conducting scientific experiments are:
- capability to install hardware on the outer surface of MRM2 for observations of
the upper hemisphere and the plane of the local horizon;
- availability of a porthole with capability to sight the hardware in the horizon
plane;
- capability to place hardware inside pressurized cabin for short periods of time for
conducting space experime nts and to store it only for the periods between EVAs;
- on-board systems resources for scientific equipment on the external surface are
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
mostly borrowed from the SM.
Fig. 4.3.2 – Characteristics of standard external workstation adapters for installing
scientific equipment units.
Location of external deliverable multi-purpose workstation on MRM2 and its
configuration are shown in Fig.4.3.3.
MRM2 resources used for integration of scientific equipment are listed in Table
4.3.1.
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SE2
SE-1
MPWS-D
Ympw
SE-3
Zmpw
Xmpw
Payload
Basepoint
2-1
Payload
Basepoint
2-2
Fig. 4.3.3 – Location of the external deliverable MPWS on MRM2 and its
configuration
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
Advanced cargo boom
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№ докум.
Подп.
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60
Инв.№ подл.
П40463
Изм. Лист
№ докум.
Подп.
Дата
Лист
61
Взам. инв. №
Инв. № дубл.
Подпись и дата
Table 4.3.1 – MRM2 resources for integration of scientific equipment
Подпись и дата
Sample fields of view of scientific equipment installed on the MRM2 BPB
multipurpose workstation are given in the following figures:
- Fig. 4.3.4 – field of view of scientific equipment installed on the PBP (axis of
sight aimed at zenith);
- Fig. 4.3.5 – field of view of scientific equipment installed on the PBP2 (axis of
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
sight aimed at zenith);
- Fig. 4.3.4 – field of view of scientific equipment installed on the PBP1 (axis of
sight aimed at zenith);
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Подпись и дата
sight aimed at zenith)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Fig. 4.3.5 – field of view of scientific equipment installed on the PBP2 (axis of
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In addition to MPW, structural elements (handrails) can also be used for scientific
equipment integration on the outer surface of MRM2.
Examples of scientific equipment installation on MRM2 handrails and fields of
view for scientific equipment are shown in Figures 4.3.6, 4.3.7.
Radiators on the USOS
Radiators on the USOS
Инв. № дубл.
Подпись и дата
FGB
Инв.№ подл.
Подпись и дата
Взам. инв. №
SM
resupply spacecraft
Fig. 4.3.6 – An example of scientific equipment installation on MRM2 handrails
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1B
2А
2B
Fig. 4.3.7 – Fields of view of scientific equipment installed on MRM2
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
1А
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4.4 Mini Research Module 1 (MRM1)
Characteristics:
- launch mass, kg: 7900;
- pressurized volume, m3: 17,4;
- storage volume for cargoes and scientific equipment, m3: 5,0;
- in particular, for scientific equipment, m3: 3.0;
- power for scientific equipment, kW: up to 0.1;
- maximum daily average rejection of heat from the hardware included in the MPF
through the air loop of the thermal control system, kW: up to 0.1;
- launch vehicle: Space Shuttle Orbiter.
MRM1 external appearance is shown in Fig. 4.4.1
FRGF
Radiative heat echanger
Docking target
Docking TV
camera
Инв. № дубл.
Подпись и дата
Airlock
Инв.№ подл.
Подпись и дата
Взам. инв. №
PVGF
ERA elbow
Active docking unit
Fig. 4.4.1 – MRM1 external appearance.
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Included in the MRM1 MPF is the mission payload equipment necessary for
providing onboard MRM1 an extended range of services in support of Russian and
commercial experiments.
Set up in the pressurized cabin are 5 multi-purpose workstations (MPWs)
equipped with mechanical adapters and mission payload equipment:
- modular shelves (up to 4 pcs.);
- Glovebox-C hardware with retractable glove box shelf;
- MBLTT thermostat;
- MBHTT thermostat;
- multipurpose vibration isolation platform MVIP.
Назначение
Slots for modular
shelves
Arch/frame
Modular shelf
frame
Подпись и дата
Pull-out shelf
Payload attachment
slots
Handle
Инв. № дубл.
Взам. инв. №
Подпись и дата
Special bolts
Velcro pads
Fig 4.4.2 – Mechanical payload adapter.
Key specifications:
- maximum payload unit dimensions 600x400x600 mm;
- maximum number of retractable shelf modules - 4
- payload attachment to shelf modules:
- anchor bolt attachment;
Инв.№ подл.
- elastic fasteners;
- Velcro fastener.
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Инв.№ подл.
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№ докум.
Подп.
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68
Подпись и дата
Инв. № дубл.
Подпись и дата
Table 4.4.1 – Characteristics of mission support equipment onboard MRM1
Взам. инв. №
Инв.№ подл.
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Взам. инв. №
Инв. № дубл.
Подпись и дата
Table 4.4.2 – MRM1 resources available for SE integration
Подпись и дата
4.5 Multipurpose laboratory module (MLM)
Characteristics:
- launch mass, kg: 20700 kg;
- pressurized volume, m3: 70.0;
- volume for scientific equipment, m3: 8,0;
- power for scientific equipment, kW: - up to 1.0 kW (daily average) inside the
pressurized cabin and up to 1.5 kW (daily average) outside the pressurized cabin;
- maximum daily average rejection of heat from the hardware included in the MPF,
kW: up to 1.0;
- launch vehicle: Proton M LV.
To support scientific equipment installation, MLM has provisions for:
- installation of payloads inside pressurized cabin of MLM (16 internal
multipurpose workstations) with the total volume of the accommodation areas of
no less than 8 m3 (including a workstation equipped with a porthole for installing
units of scientific equipment);
Взам. инв. №
Инв. № дубл.
Подпись и дата
- installation of no less than 12 simultaneous payloads outside the MLM
pressurized cabin;
- mechanical interfaces;
- vacuum interfaces;
- interfaces to provide thermostatic control for scientific equipment;
- interfaces to supply scientific equipment with power from the MLM On-board
Equipment Control System;
- command and data interfaces between scientific equipment and the MLM Onboard Equipment Control System;
- payload operating conditions in required modes when conducting space
Инв.№ подл.
Подпись и дата
experiments;
- storing original materials and obtained results.
In addition to this, to support a long-term program of scientific and applied
research, MLM provides the following equipment:
- vibration isolation platform;
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- glove box;
- thermostats.
In order to minimize EVAs when installing scientific equipment on external
workstations, the following robotic equipment is provided:
- ERA robotic arm;
- automatic airlock (AL).
The MLM external appearance is shown in Figures 4.5.1 and 4.5.2.
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
External view of the airlock is shown in Fig. 4.5.3.
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71
Инв.№ подл.
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72
Взам. инв. №
Инв. № дубл.
a cupola with a porthole
MPW-N7
Подпись и дата
Fig. 4.5.1 – MLM external appearance
MPW-N3 (Passive restraint 3)
MPW-N9
MPW-N5
Подпись и дата
MPW-N4 (Passive restraint 4)
Airlock
MPW-N8
Lab Base Point 1
(for ERA)
Инв.№ подл.
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73
Инв. № дубл.
MPW-N8
Lab Base Point 2
(for ERA)
Взам. инв. №
Подпись и дата
Fig. 4.5.2 – MLM external appearance
MPW-N4
MPW-N1 (Passive restraint 1)
Подпись и дата
Airlock
MPW-N6
MPW-N2 (Passive restraint 2)
ERA
Lab Base Point 3
(for ERA)
For installing scientific equipment outside the MLM pressurized cabin, the
following External Multi-Purpose Workstations (EMPWS) are provided:
- EMPWS1, EMPWS2 based on a passive fastener 77КМЛ.810Ю2600-0 located
outside the MLM instrumentation and cargo compartment-3 along Plane I.
Installation of scientific equipment onto EMPWS1, EMPWS2 is effected by means
of payload adapter 77КМЛ.600Ю2200-0;
- EMPWS3 based on a passive fastener 77КМЛ.810Ю2600-0 located outside the
MLM instrumentation and cargo compartment-3 between Plane II and III.
Installation of scientific equipment onto EMPWS3 is effected by means of payload
adapter 77КМЛ.600Ю2200-0;
- EMPWS4 based on a passive fastener 77КМЛ.810Ю2600-0 located outside the
airlock. Installation of scientific equipment onto EMPWS4 is effected by means of
payload adapter 77КМЛ.600Ю2200-0;
- EMPWS5 based on a passive base point 27КСМ.152Ю7200-0, located outside
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
the MLM instrumentation and cargo compartment-3 along Plane III. Used for
installing scientific equipment onto EMPWS5 is a mating device ABP
27КСМ.152Ю7110-0. It is possible to install onto EMPWS5 a multipurpose
platform
17КС.600Ю1501А-210 with active adapters 17КС.600Ю1151А-0, allowing
simultaneous installation of three payloads. Used for installing scientific
equipment onto the platform are passive payload adapters 17КС.600Ю1152А-0;
- EMPWS6 based on a passive base point 27КСМ.152Ю7200-0, located outside
the MLM instrumentation and cargo compartment-3 along Plane I. Used for
installing scientific equipment onto EMPWS6 is a mating device ABP
27КСМ.152Ю7110-0. It is possible to install onto EMPWS6 a multipurpose
platform
17КС.600Ю1501А-210 with active adapters 17КС.600Ю1151А-0, allowing
simultaneous installation of three payloads. Used for installing scientific
equipment onto the platform are passive payload adapters 17КС.600Ю1152А-0;
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- EMPWS7 based on a support 77КМЛ.600Ю1006-100 located outside the MLM
pressurized adapter between Plane III and IV. Installation of scientific equipment
onto EMPWS7 is effected by means of a base for installing the scientific
equipment onto a support 77КМЛ.600Ю1006-200;
- EMPWS8 based on a support 77КМЛ.600Ю1006-100 located outside the MLM
instrumentation and cargo compartment 3 along Plane II. Installation of scientific
equipment onto EMPWS8 is effected by means of a base for installing the
scientific equipment onto a support 77КМЛ.600Ю1006-200;
- EMPWS9 based on a support 77КМЛ.600Ю1006-100 located outside the MLM
instrumentation and cargo compartment 3 along Plane IV. Installation of scientific
equipment onto EMPWS9 is effected by means of a base for installing the
scientific equipment onto a support 77КМЛ.600Ю1006-200;
- EMPWS based on a passive fastener 77КМЛ.820Ю7360-0 located on the in-andout table of the airlock. Installation of scientific equipment onto EMPWS4 is
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
effected by means of payload adapter 77КМЛ.600Ю2200-0. The in-and-out table
of the airlock also provides structural elements for securing payloads when
conducting experiments in the extended position of the table.
Sample fields of view of scientific equipment installed on the MLM multipurpose
workstation are given in the following figures:
- Fig. 4.5.3 – field of view of scientific equipment installed on the EMPWS1 (axis
of sight pointed towards nadir);
- Fig. 4.5.4 – field of view of scientific equipment installed on the EMPWS3 (axis
of sight aimed at zenith);
- Fig. 4.5.5 – field of view of scientific equipment installed on the EMPWS3 (axis
of sight pointed towards nadir);
- Fig. 4.5.6 – field of view of scientific equipment installed on the EMPWS4 (axis
of sight aimed at zenith);
- Fig. 4.5.7 – field of view of scientific equipment installed on the EMPWS5 (axis
of sight pointed towards nadir).
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- Fig. 4.5.8 – field of view of scientific equipment installed on the EMPWS6 (axis
of sight pointed towards nadir);
- Fig. 4.5.9 – field of view of scientific equipment installed on the EMPWS7 (axis
of sight aimed at zenith);
- Fig. 4.5.10 – field of view of scientific equipment installed on the EMPWS8 (axis
of sight pointed towards nadir);
- Fig. 4.5.11 – field of view of scientific equipment installed on the EMPWS9 (axis
Подпись и дата
Взам. инв. №
Инв. № дубл.
Подпись и дата
of sight pointed towards zenith);
Fig. 4.5.3 – field of view of scientific equipment installed on the EMPWS1 (axis of
Инв.№ подл.
sight pointed towards nadir)
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Подпись и дата
sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Fig. 4.5.4 – field of view of scientific equipment installed on the EMPWS2 (axis of
П40463
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77
Инв. № дубл.
Подпись и дата
sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Fig. 4.5.5 – field of view of scientific equipment installed on the EMPWS3 (axis of
П40463
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78
Инв. № дубл.
Подпись и дата
sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Fig. 4.5.6 – field of view of scientific equipment installed on the EMPWS4 (axis of
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Инв. № дубл.
Подпись и дата
sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Fig. 4.5.7 – field of view of scientific equipment installed on the EMPWS5 (axis of
П40463
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80
Инв. № дубл.
Подпись и дата
sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Fig. 4.5.8 – field of view of scientific equipment installed on the EMPWS6 (axis of
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81
Инв. № дубл.
Подпись и дата
sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Fig. 4.5.9 – field of view of scientific equipment installed on the EMPWS7 (axis of
П40463
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Подпись и дата
of sight pointed towards nadir)
Инв.№ подл.
Подпись и дата
Взам. инв. №
Инв. № дубл.
Fig. 4.5.10 – field of view of scientific equipment installed on the EMPWS8 (axis
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of sight pointed towards nadir)
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Fig. 4.5.11 – field of view of scientific equipment installed on the EMPWS9 (axis
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Automatic airlock is intended for:
- removing payloads from the MLM pressurized adapter and placing them on the
outer surface of the space station;
- receiving payloads from ERA robotic arm and moving them into the internal
volume of the airlock and on to the MLM pressurized adapter;
- conducting scientific experiments in the internal volume of the airlock;
- conducting scientific experiments outside the airlock on the extended table and at
a specially prepared station.
Key specifications of the airlock:
- airlock mass, kg: up to 1050;
- maximum power consumption with a payload, kW: 1.5;
- gas volume, m3: 2,1;
- maximum payload mass on the payload handling device, kg: 150;
- maximum payload dimensions, mm: 1200x500x500;
- the number of vacuum pumping cycles during flight tests,
- allowable internal pressure, kg/cm2: 1.- 3;
- residual pressure after vacuum pumping, mm Hg: 10-4;
External view of the airlock is shown in Fig. 4.5.3.
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no less than: 200;
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Passive restraint
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Supports
Operator pad
the body
Modified passive
docking unit
FRGF grappling fixture
Payload moving assembly
Handrails
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Fig. 4.5.3 – Airlock external appearance
External cover assembly
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Table 4.5.2 – Characterisitics of external multi-purpose workstation adapters
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5 Cargo certification
One of the essential conditions for integrating scientific equipment into Russian
spacecraft and ISS RS is to certify the scientific equipment.
All the scientific equipment hardware delivered by Russian transportation
spacecraft to ISS RS shall be certified for transportation/functioning onboard the
spacecraft and use onboard ISS RS, that is, the hardware shall comply with the
current requirements levied upon the hardware by the Russian spacecraft and ISS
RS.
All the activities on the implementation and verification of compliance with the
current requirements for the scientific equipment and on proving its
flightworthiness are subdivided into the following groups, which correspond to the
step-by-step process of development, verification and certification of the scientific
equipment:
b) verification of the scientific equipment to check and confirm that requirements
specified for the equipment have been met. Verification includes tests,
calculations, analyses, etc. Requirements for hardware are given in paragraph 5.5.;
compliance with the specified requirements; evaluation of hardware flight
A Russian organization develops scientific equipment in accordance with
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c) analysis of verification results: verification of equipment characteristics
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a) development of scientific equipment to meet the specified requirements;
readiness; development and publication of certification documents.
requirements of НА-99, a foreign organization follows contract requirements for
conducting the space experiment.
Scientific equipment certification is conducted in order to verify its following
integral properties:
- safety (strength, fire safety, electrical safety, toxicological safety, radiation safety,
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- equipment compatibility with spacecraft and ISS RS systems;
- equipment compliance with requirements in the Specifications.
Depending on the specific kind of equipment, certification is performed as
follows:
- safety and compliance with the specified requirements are verified for all types of
equipment;
- equipment compatibility with the spacecraft and ISS RS systems are verified for
the equipment that has interfaces (mechanical, electrical, software, telemetry) with
the spacecraft and operating on-board the spacecraft.
5.1 Documents used in the course of certification
5.1.1 Documents used in certification of scientific equipment are defined in
Russian technical standards documents used for the development of rocket and
space technology products. Certification documents attesting to certification of an
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individual piece of equipment are:
- logbook, data sheet, label (hereinafter referred to as logbook) included in the
accompanying documents;
- flightworthiness statement;
- safety certificate.
5.1.2 Logbook
The logbook is intended for identifying a piece of equipment, recording its key
characteristics, and contains a flightworthiness statement for that particular unit.
The logbook is published for each piece of equipment by its manufacturer
(vendor)
in
accordance
with
the
current
requirements
for
shop-floor
documentation. In some particular cases (for example, for certain foreign-made
equipment, purchased items), when there is no logbook meeting the established
requirements, the organizational unit in charge of that item publishes a document
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functionally similar to a logbook.
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5.1.3 Flightworthiness statement.
Flightworthiness statement is intended as an assertion by the developer
(supervisor) company (organizational unit) that the current requirements of the
Specification or a document which replaces it (supplement to a contract, etc). The
Statement shall attest to the completion of the scheduled developmental testing, to
the fulfillment of the current requirements for scientific equipment, and contain a
permission to fly the equipment. The Statement is the primary certification
document published by the developer based on the results of a series of activities
aimed at analyzing, developmental testing and confirmation of the scientific
equipment readiness for flight/flight testing per requirements of Russian technical
standards.
The Statement is published for the scientific equipment for which a permission to
fly is sought for the first time (newly developed or structurally modified
equipment), and which is used under the Russian program of scientific and applied
research.
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The Statement’s scope shall cover the flight of scientific equipment inside a
particular type of a transportation spacecraft and its use within an ISS RS module.
5.1.4 Safety Certificate
The Certificate attests to the scientific equipment compliance with the current
safety requirements. The Certificate is published based on the results of analysis of
the equipment safety characteristics and measures taken to prevent hazardous
situations. The results of such an analysis are presented in the form of a report
(attachment to the certificate), which, together with the certificate constitute a
single Safety Data Package (SDP).
The safety certificate is published:
- for scientific equipment of foreign organizations: for the phase of its
transportation on a Russian spacecraft and for the phase of its use and storage onboard ISS RS;
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- for Russian scientific equipment: for the phase when it is used and stored onboard
ISS RS, and the phase when it is transported onboard a foreign spacecraft.
If the scientific equipment is a re-flight item for a particular phase (a certificate
was published for an earlier successfully completed flight), safety certificate is not
published for that item for that particular phase.
5.2 Scientific equipment tests
5.2.1 Generally, scientific equipment is subjected to the following kinds of tests:
- qualification tests (QT);
- acceptance tests (AT);
- tests at checkout and testing facility within the flight model of the spacecraft/ISS
RS module;
- incoming inspection at the processing facility;
- tests at the processing facility within the flight model of the spacecraft/ISS RS
All the tests and operations to service and process the scientific equipment at
RSC Energia and the processing facility are only conducted per RSC Energiaapproved instructions and under supervision of its responsible representatives. The
equipment.
qualification model of the equipment, which is not intended for operation, but
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instructions shall contain all the information that is necessary to handle the
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module;
5.2.2 Qualification tests are conducted by the equipment developer on the
which is built per documentation for the flight model of the equipment.
The prime task of qualification tests is to check the equipment against
specification or contract requirements under exposure to environments which are
as close to operational environments as possible, and, if need be, under heavier
duty test modes.
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5.2.3 Acceptance tests are done to check the equipment against the requirements of
the documentation approved by RSC Energia. The procedure for conducting
acceptance tests is determined by the contract requirements.
5.2.4 Tests at the checkout and testing facility within the flight model of the
spacecraft/ISS RS module are conducted in order to check mechanical, electrical
and software interfaces between the equipment and the spacecraft/module.
Interfaces with the spacecraft are checked in the course of spacecraft processing
at the checkout and testing facility; interfaces with an ISS RS module are checked
in the course of joint tests of scientific equipment with the electrical model of the
module at the checkout and testing facility.
Tests within the flight model of the spacecraft at checkout and testing facility and
at the processing facility are conducted when interfaces are available by agreement
with RSC Energia.
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5.2.5 Incoming inspection of the equipment at the processing facility is conducted
per agreed procedure after its transportation to the processing facility.
Incoming inspection operations include:
- inspection of external appearance and a check for missing items;
- equipment operational integrity check (if need be).
5.2.6 Program of qualification and acceptance tests
Requirements for equipment tests are defined in ОСТ 92-5100-89 and SSP
50094.
A typical list of tests and verification activities under the programs of
qualification tests and acceptance tests of the equipment is given in Table 5.2.1.
For each type of tests a test procedure is developed.
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Table 5.2.1
Test type
№ докум.
Verified requirements
Test/verification
Подп.
1. Check for missing items
(subparagraph # in
Qual.
Section 6)
test
Requirements of design –
+
Checks delivery set for missing items and
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documentation for the
completeness of technical and operational
scientific equipment
documentation
2. Inspection of the external Requirements of design –
appearance
Note
AT
+
The external appearance of the equipment is
documentation for the
checked against outline drawing, and it is
scientific equipment
checked whether logbooks and certificates are
available and completely filled out
3. Dimensions check
Requirements of design –
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hardware,
scientific equipment
mounting holes, mounting holes dimensions
+
documentation for the
Requirements of design –
documentation for the
scientific equipment
center-to-center
distances
+
of
Mass and center of mass position are checked
against the developed outline drawings
scientific equipment
5. Packing check
To be checked are: outline dimensions of the
documentation for the
4. Checking mass and position Requirements of design –
of the center of gravity
+
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Verified requirements
№ докум.
Test/verification
6. Safety review
Test type
(subparagraph # in
Qual.
Section 6)
test
6.3
–
Note
AT
+
Safety analysis includes safety assessment in
Подп.
general
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and
specialized
fields
(strength,
safety, etc.), as well as an integrated assessment
materials safety, toxicological and radiation
of the hardware safety,
7 Mechanical loads tests
The tests are performed to check the strength of
7.1. Vibration
+
+
load-bearing elements of the hardware that has
7.2. Shock loads
+
–
a mechanical interface (rigid attachment) with
7.3. Linear accelerations
р
–
the spacecraft/module, as well as, if need be, to
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check the functional integrity of the hardware.
6.1.3
for
spacecraft
6.2.2 for ISS RS
The tests shall be performed in transportation
and working configurations of the hardware
under the types of loading which correspond to
these configurations (for example, in a
bag/package fastened in a way similar to the
standard attachment to the spacecraft).
If the hardware was tested at levels different
from those currently established, the issue of
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Verified requirements
№ докум.
Test/verification
Test type
(subparagraph # in
Qual.
Section 6)
test
AT
Note
extending the applicability of these results
Подп.
needs to be coordinated with RSC Energia.
8 Environmental tests
Дата
8.1. Humidity
6.1.2.3 for spacecraft,
+
–
be agreed with RSC Energia
6.2.1.5 for ISS RS
8.2
Temperature
The accuracy of humidity measurements is to
when 6.1.2.2
+
-
6.1.2.2 for spacecraft +
+
delivered to the processing
facility
8.3 Temperature in flight
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at lower/higher temperature
6.2.1.4 for ISS RS
8.4 Thermal cycle
Requirements
The tests (for thermal stability) are conducted
of +
–
Subjected to the tests is scientific equipment
specifications for the
operating in unpressurized compartments under
scientific equipment
exposure to cyclically varying temperature
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Verified requirements
№ докум.
Test/verification
8.5 Acoustic noise
Test type
(subparagraph # in
Qual.
Section 6)
test
6.1.3.3 for spacecraft +
Note
AT
–
Подп.
6.2.3.3 for ISS RS
The tests are conducted if the equipment is
considered critical with respect to acoustic
noise exposure, as well as in the cases where
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the root-mean-square value of the acoustic
pressure on the hardware in the locations where
it is installed in the spacecraft/module exceeds
130 dB
8.6 High/low pressure tests
6.1.2.4 for spacecraft +
-
The tests are conducted in order to check
operational integrity of the equipment under
6.2.1.3 for ISS RS
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exposure to extreme pressure values
8.7 Pressure drop rate tests
6.1.2.4 for spacecraft
+
-
Subjected to the tests is scientific equipment
built in the form of a pressurized unit or a
vented container used inside the descent
vehicle or orbital module of Soyuz spacecraft.
Pressure drop rate is 100 mm Hg/s
9 Radiation stability tests
6.1.2.5 for spacecraft +
6.2.6 for ISS RS
–
The test is run for the kinds of equipment that
are radiation exposure critical
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Verified requirements
№ докум.
Test/verification
10 Leakage tests
Test type
(subparagraph # in
Qual.
Section 6)
test
Requirements of
+
Note
AT
+
Subjected to the tests is the equipment that has
Подп.
specifications for the
a pressurized volume (an enclosure, a vessel
scientific equipment
containing gas, liquid, test tubes with biological
Дата
samples, etc.).
Before and after strength tests the equipment
must be tested for leakage.
Before filling, the flight unit must be tested for
leakage.
11 Thermal vacuum test
6.1.2.2 for spacecraft +
–
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6.2.1.4 for ISS RS
Subjected to tests is the equipment that
operates in open space in unpressurized
compartments. It is recommended that the tests
be combined with thermal stability tests (see
paragraph 8.3)
12. A check of electric circuit
Requirements of
specifications and
design documentation
for the scientific
equipment
+
+
A check of electrical circuits continuity and
isolation.
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Verified requirements
№ докум.
Test/verification
(subparagraph # in
Qual.
Section 6)
test
13 Insulation and insulation
Подп.
breakdown tests
6.1.4 for spacecraft,
14. Inrush current test
6.2.4 for ISS RS
Дата
15. Power consumption check
16. Functional test
Test type
Requirements of
Note
AT
+
+
–
+
–
+
+
+
Acceptance tests are performed at nominal
specifications for the
supply voltage, while qualification tests are
scientific equipment
performed at those extreme voltages to which
the scientific equipment is most sensitive under
given conditions
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17. EMC test
6.1.2.5 for spacecraft
6.2.6 for ISS RS
+
–
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Verified requirements
№ докум.
Test/verification
Test type
(subparagraph # in
Qual.
Section 6)
test
Note
AT
All flight units of self-contained power sources,
Подп.
except the ones of the “tablet” type are subject
Дата
recording
to tests. The tests include: visual inspection,
physical
characteristics,
taking
measurements of storage voltage and operating
voltage,
leakage
test,
cycling
18. Tests of self-contained
(charge/discharge).
power supplies and chargers
Tablet-type self-contained power sources are
tested within operating equipment.
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Flight units of chargers are tested under normal
operating conditions.
Self-contained power sources and chargers that
have passed the tests within the equipment are
not subjected to tests within this scope.
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Notes:
№ докум.
1) Symbols:
- "+" – this test must be conducted;
- "р" – these requirements can be verified through calculation (analysis);
Подп.
- tests/checks are not conducted
Дата
2) It is acceptable, with the agreement of RSC Energia, not to conduct certain types of tests specified in the table. In that
case, the equipment compliance with the requirements shall be verified through analysis.
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6 Specifications for experimental equipment delivered to ISS
6.1 Specifications for equipment transported in Russian spacecraft Progress
and Soyuz
6.1.1 General
These requirements are applicable to all the cargoes delivered to ISS in Progress
and Soyuz spacecraft, to all the waste and cargoes removed from the station for
disposal inside Progress cargo compartment or Soyuz orbital module, as well as to
cargoes returned to Earth inside Soyuz descent vehicle. The document contains
requirements for all the phases of transportation and for the spacecraft ground
processing phase.
All the transported cargoes shall meet the requirements of the normal operating
conditions described in this Section. In off-nominal situations, including
depressurization, cargoes shall remain fire-and-explosion safe, nontoxic, leakproof
(gases, liquids, biological samples) and preserve its mechanical integrity, meet
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safety requirements for crew and spacecraft equipment. There are no requirements
to assure cargo functioning after an off-nominal situation.
In accordance with the conditions required onboard spacecraft for cargo delivery,
all the cargoes are subdivided into:
- active cargoes, which require for their transportation connections to the
spacecraft onboard systems (power supply system, thermal control system,
telemetry, etc.);
- passive cargoes, which do not require for their transportation any connections to
the spacecraft onboard systems.
6.1.2 Environmental conditions
6.1.2.1 Gas medium
Composition of the gas medium inside Cargo Compartment, Descent Vehicle
and Orbital Module (in volume percentages) through all the phases of free flight up
to the docking with the space station:
- oxygen − 21...40% (at partial pressure of 120...350 mm Hg);
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- carbon dioxide – up to 3%;
- hydrogen – up to 2%;
- helium – up to 0.01%;
- nitrogen – the rest.
Gas media composition in Cargo Compartment, Descent Vehicle and Orbital
Module after docking with the space station and opening of the transfer hatches is
the same as the composition of the atmosphere inside the space station pressurized
cabin.
Phase of the operation
Temperature
In production and storage rooms
+5÷+30°C
Ground transportation
±50°C
Assembly and testing
0÷+40°C
On the launch pad (for cargoes)
0÷+40°C
Orbital flight
- for cargoes, when the crew is not present
0÷+40°C
- for cargoes, when the crew is present
+18÷+25°C*)
- for cargoes outside pressurized compartments
from minus 150 to +
125оС
After spacecraft landing (for cargoes)
*)
±50°C
− Temperature is allowed to rise up to +30°C and drop down to +10° for no more
than 3 hours a day.
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6.1.2.2 Temperature
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6.1.2.3 Humidity
(at temperature +20°C)
Phase of operation
Relative humidity
In production and storage rooms
20...85%
Transportation
up to 90%
Assembly and testing
20...80%
On the launch pad (for cargoes)
20...85%
Orbital flight (for cargoes, when the crew is not present)
20...80%*)
Orbital flight (for cargoes, when the crew is present)
30...75%
After descent vehicle landing (for cargoes)
up to 98%
*)
− A rise up to 90% for as long as 3 hours a day is possible.
6.1.2.4 Pressure
Absolute pressure of the gas medium inside pressurized compartments through
mission phases it may vary within 450...970 mm Hg.
The vacuum outside the spacecraft (in the cargo compartment of Progress
spacecraft when conducting experiments requiring its depressurization, or when
cargoes are located on the outer surface of the spacecraft) is at 10-3...10−8 mm Hg.
In case of depressurization of the cargo compartment or the descent vehicle of
Soyuz, the cargo may stay at a pressure as low as 10-6 mm Hg for as long as 10
hours. The maximum pressure drop rate is 100 mm Hg/s.
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all the phases of ground processing is the atmospheric pressure, and through all the
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6.1.2.5 Radiation exposure
Radiation exposure requirements for Progress and Soyuz spacecraft are defined
in SSP 50094, Subsection 3.6.
To evaluate exposure to external radiation sources, the following values are
assumed to characterize the protection: during Progress missions, for elements
inside the compartment located near to the hull – 1 g/cm2, during Soyuz missions,
inside compartments near the hull − 1g/cm2, and under the couches in the descent
vehicle − 3 g/cm2.
Radiation from the internal source inside the descent vehicle is a flux of gammaray photons with energies of up to 9.7 MeV. The dose rate does not exceed 0.1
rad/day at the distance of 0.6 meters from the source (the area under the couches).
6.1.3 Mechanical loads
6.1.3.1 Coordinate systems
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Ground transportation:
X1 axis is parallel to the direction of transportation. Positive direction of the axis is
against the transportation direction;
Y1 axis is perpendicular to X1 axis. The positive direction of the axis is up;
Z1 axis is perpendicular to axes X1, Y1 and completes the right-handed coordinate
system.
Orbital insertion and flight:
X axis is the longitudinal axis of the vehicle. Positive direction is towards
maximum loads during ascent (Attachment 2);
Y axis is perpendicular to X axis;
Z axis is perpendicular to axes X, Y and completes the right-handed coordinate
system.
The XYZ coordinate system is rotated around X axis by 135° with respect to X1 Y1
Z1 coordinate system.
Descent vehicle reentry and landing:
X axis – The positive direction is towards maximum loads;
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Y axis is perpendicular to X axis. The positive direction is towards maximum
loads;
Z axis is perpendicular to axes X, Y and completes the right-handed coordinate
system.
6.1.3.2 Types of mechanical loads
All types of loads can be grouped as follows:
- shock loads of short duration occurring during transportation, launch vehicle and
spacecraft engine firings and shutdowns, separation of the spacecraft from the
launch vehicle, spacecraft separation during reentry, etc.;
- vibration loads produced by acoustic and mechanical effects during operation of
propulsion systems and atmospheric turbulence;
- linear and low-frequency dynamic loads.
Cargoes that are highly susceptible to shock and vibration loads shall be installed
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on shock absorbers by the developer.
6.1.3.3 Loads
6.1.3.3.1 When the spacecraft is transported within the processing facility
When the spacecraft is transported by rail within the processing facility, the
operational loads are:
- nх1= ±1.5; ny1= –1 ±0.25; nz1= ± 0.2;
When lifted, carried and installed onto supports:
− in horizontal position:
nх1= ±0.1; ny1= – 1±0.5; nz1= ±0.1;
− in vertical position:
nх1= –1.5; ny1= ±0.1; nz1= ±0.1.
6.1.3.3.2 During ascent to orbit
Operational values for linear loads for cargoes (without taking into account the
low-frequency dynamic component) are:
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- along X axis – +4.3;
- along Y,Z axes – ±1.5.
Operating values for quasi-static loads (taking into account the low-frequency
dynamic component) for payloads secured in transportation position, which do not
have (with the effects of transportation fixtures taken into account) any normal
modes within the frequency range of up to 30 Hz, or have normal modes in the said
frequency range, while having a Q-factor for the cargo-fasteners system of no more
than 10, are 7.15 in any of the three mutually perpendicular directions occurring
non-simultaneously.
Duration of exposure to linear and quasi-static loads: up to 600 s.
Vibration loads in 3 mutually perpendicular directions in the form of sine vibration
are given in Table 6, and, at frequencies above 20 Hz in the form of random
vibration, are given in Table 5.
Shock loads are in accordance with Table 1.
Operational acoustic loads on the spacecraft:
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- acoustic noise levels (total acoustic pressure) during ascent to orbit are: outside
spacecraft – up to 145 dB.
The acoustic noise spectrum inside the payload fairing of the launch vehicle is
given in Table 7.
Duration of exposure: 20÷30 s.
- inside cargo compartment and the orbital module – up to 135 dB.
The acoustic noise spectrum is given in Table 8. Qualification levels are presented,
duration of exposure is 120 s.
inside descent vehicle – up to 125 dB.
6.1.3.3.3 When Soyuz emergency rescue system (escape tower) operates (offnominal situation)
Operational values for linear loads:
a) during descent vehicle escape manuever:
nх=+15, exposure duration is 5 s,
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nх=-8, exposure duration is 2 s,
nz,y=± 13, 1 Hz vibration, t=30 s;
b) during ballistic descent of the descent vehicle after launch vehicle failure:
nх=+27, exposure duration is 9 s,
ny=+5, exposure duration is 13 s,
nz=±1, exposure duration is 6 s.
Vibration loads in three mutually perpendicular directions, applied to the
escaping part as random vibration per Table 5.
Shock loads are in accordance with Table 4.
Operational acoustic loads on the spacecraft
Acoustic noise levels (total acoustic pressure) during operation of the emergency
rescue system (operation time is 4 s) are:
- outside the spacecraft – up to 160 dB;
- inside the orbital module – up to 150 dB;
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- inside the descent vehicle – up to 140 dB.
6.1.3.3.4 In orbital flight:
1) operating values for quasi-static loads (taking into account the low-frequency
dynamic component) for payloads secured in transportation position, which do not
have (with the effects of transportation fixtures taken into account) any normal
modes within the frequency range of up to 30 Hz, or have normal modes in the said
frequency range, while having a Q-factor for the cargo-fasteners system of no more
than 10, are in longitudinal direction: nx=±0.3, in lateral direction: nx=±1.2, in
transverse direction: nz=±1.2 (simultaneous exposure) for 12 minutes;
2) vibration loads for all spacecraft compartments in 3 mutually perpendicular
directions in the form of sine vibration are given in Table 6.1.3.3.5.3, and, at
frequencies above 20 Hz in the form of random vibration, are given in Table
6.1.3.3.5.2;
3) shock loads are given in Table 6.1.3.3.4.2;
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4) operational loads for dynamic loads during docking with ISS are given in Table
6.1.3.3.4.3.
Table 6.1.3.3.4.1
In the X-axis direction
In two other mutually Pulse duration,
perpendicular directions ms
Number of Load
Number of
impacts
impacts
2
40
2
1-3
Load
Descent
40
vehicle, orbital
module
and
cargo
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compartment
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Table 6.1.3.3.4.2
Loads in the X-axis direction
Loads in two other mutually
perpendicular directions
Spacecraft
Accelerat Numb
Pulse
Accelerat Numbe Pulse duration,
compartment ion, (g) er of duration, (ms) ion, (g)
r of
(ms)
s
impac
impacts
ts
Descent
vehicle,
40
5
1-3
40
5
1-3
orbital
module
and
cargo
compartment
Table 6.1.3.3.4.3
In two other mutually perpendicular
directions
The number
Cycle
Load
The number
Cycle
of loading
duration,
of loading duration, (ms)
cycles
(ms)
cycles
≤10
100÷1000 ±1,0±0,2
≤10
100÷1000
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Loads in the X-axis direction
Load
±0,8
6.1.3.3.5 During Soyuz descent and landing
1) operational values for linear accelerations:
a) normal descent of the descent vehicle (nominal conditions)
nх= 0…+6, ny,z= 0±1, duration – 600 s, including
nх=+6, exposure duration is 20 s;
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ny,z= ±1, exposure duration is 20 s;
a) ballistic descent of the descent vehicle (off-nominal situation):
nх= 0…+10, ny,z= 0±1 for 600 s, including
nx= +10, exposure duration is 20 s;
ny,z= ±1, exposure duration is 20 s;
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Variation of operating values of loads with time is approximated by a sine halfwave with half-period equal to the exposure duration;
c) during parachute system operation:
- during drogue chute operation
n= +10 (in any direction), exposure duration is 10 s;
- during main chute operation
nx= +8, exposure duration is 15 s;
ny,z= ±4, exposure duration is 15 s;
- during re-hooking
n= +5 (in any direction), exposure duration is 5 s;
d) touch-down with failed soft-landing thrusters (off-nominal situation):
at first impact nх= +100 (a deviation of loading direction of ±15° is possible),
ny,z=50 (direction of loads in transverse plane is arbitrary).
Variation of loads with time is approximated by a sine half-wave with halfperiod τ=20-30 ms;
vehicle during reentry in the form of random and sine vibration in accordance with
Tables 6.1.3.3.5.2, 6.1.3.3.5.3:
3) shock loads are in accordance with Table 6.1.3.3.5.1.
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2) vibration loads (in three mutually perpendicular directions) on the descent
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Table 6.1.3.3.5.1
Loads in two other mutually
perpendicular directions
Accelera Number
Pulse Accelera Number
Pulse
tion, (g)
of
duration, tion, (g)
of
duration,
impacts
(ms)
impacts
(ms)
100
2
1-2
100
2
1-2
Loads in the X-axis direction
Spacecraft
compartment
Descent vehicle
(during escape tower
40
3
3-5
40
3
3-5
Descent vehicle
40
2
1-3
40
2
1-3
during cargo return
20
7
1-3
20
7
1-3
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operation)
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Table 6.1.3.3.5.2
Spacecraf
Mission
t
phases
compart
Exposure
Frequency subranges, Hz
duration, s
ment
20-50 50100
100-
200-
500-
1000-
200
500
1000
2000
Vibration acceleration spectral density, g2/Hz
Ascent to
Orbital
orbit
module,
0,02
0,02
0,02-
0,05
0,05
0,05-
0,025-
0,025
0,013
0,02-
0,01-
0,01
0,005
120
cargo
compart
ment
Descent
0,02
0,02
0,02
0,02
vehicle
Подпись и дата
Orbital
0,02
0,02-
0,008- 0,004-
0,008
0,004
0,002
0,004
0,004
0,004-
480
cargo
compart
ment,
Инв. № дубл.
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0,02
module,
descent
vehicle
Orbital
Orbital
flight
module,
0,004
0,004
0,004
600
0,002
cargo
compart
Подпись и дата
Инв.№ подл.
0,02
120
ment
Descent
0,004
vehicle
0,004
0,004
0,004- 0,01
0,01-
0,01
0,005
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Spacecraf
Mission
t
phases
compart
Exposure
Frequency subranges, Hz
duration, s
ment
20-50 50100
100-
200-
500-
1000-
200
500
1000
2000
Vibration acceleration spectral density, g2/Hz
Descent
Descent
0,004
0,004
0,004
vehicle
Escape
Descent
0,08-
0,2-
tower
vehicle
0,2
0,4
0,4
0,004- 0,01
0,01-
0,01
0,005
0,4-
0,16-
0,08-
0,16
0,08
0,04
600
10
operation
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Loads vary linearly with frequency within a subrange.
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Table 6.1.3.3.5.3
For all
Exposure
Frequency subrange, Hz
spacecraft
duration, s
compartments
Mission phases
5-25
25-200
200-800
800-1500 15002500
Vibration acceleration amplitude, g
Ascent to orbit
1
1-3
3-5
5-8
8
300
Orbital flight
0,5
0,5 - 1
1-3
3-5
5-2
300
Descent
−
0,5 - 1
1-3
3
3-2
100
2 - 10
10 - 15
15 - 10
10 - 5
10
Escape
tower 1 - 2
operation
Loads vary linearly with frequency within a subrange.
Table 6.1.3.3.5.4
31,5
63
125
250
500
1000
2000
4000
120
118
2000
-
100
-
Maximum root-mean-square level of acoustic pressure, dB
129
134
138,5
136
135
127
Table 6.1.3.3.5.5
Center frequency of the octave frequency subrange, Hz
31,5
63
125
250
500
1000
Maximum root-mean-square level of acoustic pressure, dB
119
119
130
131
129
119
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Center frequency of the octave frequency subrange, Hz
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6.1.3.3.6 Autonomous transportation of equipment and hardware
For autonomous transportation of equipment and hardware by all types of
transportation, the recommended accelerated test modes are given in Table
6.1.3.3.6.
Table 6.1.3.3.6
The number
Shock
pulse
Transportation vehicle
Pulse duration,
acceleratio
axes
ms
n, g
Total
of impacts
number of
per minute
impacts
(no more
than)
X1
Y1
Z1
Number of impacts
9
750
2500
1750
5 to 10
5000
120
Strength analysis of deliverable cargoes, including their fasteners, is done based
on design overloads defined as the product of operating overloads and the safety
factor f.
The same safety factors f are also taken into account during qualification tests.
Minimal safety factors are set in accordance with Table 6.1.3.4. When the
structure is exposed to a combination of forces for the loads increasing the strength
margin, the safety factor is set at f=1. Loads increasing the strength margin are
determined during strength analysis.
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6.1.3.4 Safety factors
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Table 6.1.3.4
Loads per paragraphs
Value of safety factor f, no less than
6.1.3.3.1
1,5
6.1.3.3.2
1,4
6.1.3.3.3
1,3
6.1.3.3.4
2,0
6.1.3.3.5) a,b,c
1,5
6.1.3.3.5) d
not specified
In the course of strength analysis an additional coefficient K is introduced, which
is a multiplier for the safety factor:
- for fasteners (for example: bolts, screws taking into account a pre-torque) К =
1.25;
- for critical fasteners having a complex geometric shape (for example: brackets,
- when using materials having a high spread in strength properties or a significant
anisotropy (for example: engineering plastics), К=1.2;
6.1.3.5. Vibration and acoustic tests
least as high as those given in paragraph 6.1.3.
vibration per Table 6.
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6.1.3.5.1 Vibration and shock qualification tests are conducted at levels that are at
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fittings), К=1.25;
6.1.3.5.2 It is allowed to replace random vibration tests per Table 5 with sine
6.1.3.5.3 Operational acoustic loads applied to the spacecraft during qualification
tests are increased by no less than 3 dB, and the exposure duration by no less than
a factor of 4, but no less than 120 s).
6.1.3.5.4 G-load qualification tests are conducted taking into account safety factors
per paragraph 6.1.3.4.
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6.1.3.6 Microgravity
Active cargoes that operate continuously or in the microgravity mode shall meet
the requirements for microgravity conditions specified in paragraph 6.2.3
6.1.4 Electrical conditions and requirements for electrical equipment
Electrical conditions and requirements are given for active cargoes which require
connection to the spacecraft onboard power system.
Active cargoes are connected to the spacecraft onboard power system via fourpin connectors to outlets rated at 3A (Progress M1), 10A and 20A using a circuit
breaker.
6.1.4.1 Onboard power system voltage
The spacecraft onboard power system provides DC power at 27 +−74 V.
6.1.4.2 Bonding to protect against static electricity
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In order to assure that all the conducting parts of the cargo and the spacecraft
have the same electrical potential, all the electrically conductive structural
elements need to be bonded, that is, have reliable electrical interconnections with
contact resistance of no more than 2.5 mOhm.
6.1.4.3 Insulation resistance
Insulation resistance of electrical circuits in active cargoes shall be no less than
20 MOhm at relative air humidity of 45...80% within the temperature range of
+15...+35°C, and no less than 1 MOhm at relative air humidity of 95% and +20°C
temperature.
Insulation of primary current-carrying circuits from the device case and between
any electrically isolated circuits shall be able to withstand a test voltage of 200 V
(effective or DC voltage) for 1 s.
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6.1.4.4 Design features
Design of high-voltage elements in active cargoes shall rule out any possibility of
corona discharges and short-circuiting at low pressure inside compartments.
All current-carrying elements in active cargoes having non-insulated electric
conductors located in the air flow shall be securely protected against contact with
small objects using airtight covers.
6.1.4.5 Noise immunity and electromagnetic compatibility
Requirements for electrical and RF equipment to provide noise immunity and
electromagnetic compatibility when exposed to inadvertent interference finding its
way into the equipment in a typical noise environment are given in paragraph
6.2.4.2..
6.1.5 Cargo labeling requirements
The cargoes to be delivered or removed are to be labeled with onboard data
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sheets and barcode labels, as well as photographed and filmed on video.
Deliverable cargoes shall have onboard data sheets and barcodes to make them
identifiable per requirements of paragraph 6.2.5.
Barcode labeling is done by affixing special labels carrying a barcode duplicated
in alphanumeric format. Barcode labels must not cover any existing inscriptions on
the cargo.
The locations for onboard data sheets and labels are determined by the developer
or a person in charge of the cargo and are specified in its dimensional installation
drawing. If the cargo is not subject to barcode labeling, a “barcode is not to be
installed” entry shall be made in the dimensional installation drawing.
Photographs of the delivered cargo are taken separately prior to installation on
the spacecraft, as well as after its installation on the spacecraft.
Deliverable cargoes are loaded into the spacecraft and checked for labeling per
design documentation requirements.
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6.1.6 Requirements for cargo mechanical properties
Mechanical properties shall be consistent with the cargo delivery and disposal
conditions. The cargo shall be able to go through the spacecraft hatches. The
cargoes to be delivered are loaded into Progress spacecraft through the hatch in the
docking assembly (inside diameter of 800 mm) and through three loading hatches
(inside diameter of 470 mm) in the cargo compartment. Inside diameter of the
boarding hatch in the orbital module of Soyuz is 660 mm, the diameter of the
manhole hatch in the Soyuz descent vehicle is 620 mm. The dimensions of the
cargoes are determined during approval of the dimensional installation drawings
for these cargoes.
Cargoes to be delivered can be secured inside Progress and Soyuz by means of:
- flanges and bolts;
- bands and straps;
- putting cargoes inside containers.
Mass, dimensions, inertia properties of the cargoes to be delivered shall conform
6.1.7 Packing requirements
The packing used in flight shall preserve cargo integrity during transportation
and storage within the transportation spacecraft.
The packing material used in flight shall comply with SSP50094 paragraph 4.3.4.
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to the outline drawings approved by the Russian side.
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6.2 Specifications for the equipment stored and operated onboard ISS RS
6.2.1 Medical and technical requirements for habitable environment
6.2.1.1 Atmospheric composition
Maximum content of major components in the space station atmosphere in orbit
(volume %):
- nitrogen: up to 78 % (no more than 600 mm Hg);
- oxygen: up to 24.8 %
- carbon dioxide: up to 3 %
- methane: up to 0.5 %
- hydrogen: up to 2.0 %
- helium: up to 0.01 %
- water vapor: up to 3 % (up to relative humidity of 90 %).
6.2.1.2. Harmful trace contaminants content
The maximum content of harmful trace contaminants shall comply with Table
acceptable risk (the duration of a period within an expedition to orbit during which
acceptable risk standards apply must be additionally defined).
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6.2.1. Russian standards correspond to zero risk, US standards correspond to
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Table 6.2.1 Maximum acceptable content of harmful trace contaminants in the
space station atmosphere
Compound 1)
Russian MAC values
2)
US SMAC values 3) for
for 360 days in orbit 180
days
mg/m3
mg/m3
hydrogen
1600
340
methane
3300
380
pentane
10
590 (7 days)
hexane
5
180 (7 days)
heptane
10
200 (7 days)
formaldehyde
0,05
0,05
acetaldehyde
1
4
aliphatic
aldehydes 1
in
orbit
4 to 8
0,02
0,03
methyl alcohol
0,2
9
ethyl alcohol
10,0
2000
2-propyl alcohol
1,5
150
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1-butyl alcohol
0,8
40
acetone
2
50
Взам. инв. №
2-butanone
0,25
30
benzole
0.2 (180 days)
0,2
toluene
8
60
xylene
5
220
sterol
0,25
43 (7 days)
isopropylbenzol
0,5
49 (7 days)
furan
0,05
0,025
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propionic aldehyde
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(benzaldehyde)
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Compound 1)
Russian MAC values
2)
US SMAC values 3) for
for 360 days in orbit 180
days
mg/m3
mg/m3
ammonia
1
7
ethyl acetate
4
-
carbon monoxide
5
10
polymethyl cyclosiloxane
0,2
9-15
dichloromethane
5
10
1,2-dichloroethane
0,5
1
Freon-218
150
85
in
orbit
1)
compounds are grouped into structural classes
2)
Russian Maximum Allowable Concentrations (MAC) are stated in the national
standard GOST Р 50804-95
3)
US Maximum Allowable Concentrations (SMAC) are stated in the document
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"Maximum Allowable Concentrations for individual contaminants in spacecraft
atmosphere" (JSC 20584)
6.2.1.3. Atmospheric pressure
The total nominal pressure onboard the space station is maintained within 734770 mm Hg, the minimal pressure is no less than 700 mm Hg.
6.2.1.4. Atmospheric temperature
Atmospheric temperature:
- in the living area - 18-28 оC;
- in the instrumentation area - 10-40 оС (during manned missions) and 0-40 оС
during unmanned missions.
Temperature outside pressurized compartments (in open space):
-150 to + 125оС
6.2.1.5. Humidity.
Relative humidity: 30-70%, up to 95% for short periods of time (up to 3 hours
per day). Dew-point temperature: 4.4-15.6 оC.
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6.2.1.6. Aerosol content.
Aerosol content in the atmosphere: no more than 0.15 mg/m3 for particles
ranging from 0.5 to 300 micron.
6.2.1.7. Medical and technical requirements for microbiological content.
Requirements for the allowable number of microorganisms are given in Table
6.2.2.
Table 6.2.2
Source
Air
Sampling time
(colony- prior to flight
Number of bacteria
Number of fungi (no more
(no more than)
than)
300
50
in flight
100
100
prior to flight
5,0
0,1
in flight
100
1,0
forming
units/m3)
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Internal
surface
(colonyforming
units/m2)
There should be present no pathogenic bacteria or fungi.
6.2.2 Mechanical loading conditions
Equipment loading and microgravity conditions and requirements given below
apply to the loads occurring during storage and operation of the equipment in orbit
within ISS RS. Requirements for other operational phases are defined separately
in accordance with the documentation currently in effect. In particular, conditions
and requirements for launching with the use of Soyuz transportation spacecraft and
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Progress logistics spacecraft, and for returning cargo using Soyuz are given in
Section 6.1 of this document.
If the equipment meets the below requirements, it can be stored and operated (as
far as mechanical loading conditions go) in any ISS RS module. If need be,
mechanical loading modes for a specific piece of equipment can be updated to take
into account its specific location on the ISS RS.
6.2.2.1 Loads
6.2.2.1.1 Vibration loading
Vibration modes are specified by means of vibration acceleration amplitudes
within the up to 20 Hz frequency range and vibration acceleration spectral density
values in the frequency range above 20 Hz. For the up to 20 Hz frequency range
the modes are listed in Table 6.2.3. Random vibration modes within the 20-2000
Hz frequency range are given in Table 6.2.4. The modes are specified for three
mutually perpendicular directions. Modes listed in Table 6.2.3 and 6.2.4. do not
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apply to the equipment located inside and outside propulsion compartments of
spacecraft, SM and other modules, when propulsion systems located inside these
compartments are operating. Mechanical loading modes for such equipment are
specified separately in accordance with the documentation currently in effect.
Table 6.2.3 Sine-wave vibration modes
Frequency range, Hz
Vibration acceleration
Time of application in
amplitude, g
each direction, s
1-5
0,2-0,4
2400
5-10
0,4-0,5
1200
10-20
0,5
1200
Notes: 1. The modes listed in the table correspond to the qualification test levels.
2. Vibration acceleration values vary linearly between frequencies.
3. The time of application of vibration acceleration in frequency subranges is
specified in the tests using the method of smooth variation of frequency.
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Table 6.2.4 Random vibration modes
Frequency, Hz
20
50
100
200
Time of
500
1000
2000
each
Vibration acceleration spectral density, g2/Hz
0,004
0,004
0,004
0,004
0,004
0,004
application in
direction, s
0,002
2700
Notes: 1. The modes listed in the table correspond to the qualification test levels.
2. The values for spectral densities between the above frequencies vary linearly with
logarithmic scale for frequency and spectral density.
If need be, the above vibration modes for an individual piece of equipment can
be updated to take into account its specific location on the ISS RS.
6.2.2.1.2 Dynamic and cyclic loading of the equipment during in-orbit storage and
operation within ISS RS
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The structure of the equipment installed inside and outside the ISS RS modules
shall be capable of withstanding inertial loads caused by forces acting on the ISS
structure (contact forces from dockings and undockings of various spacecraft,
forces generated by operation of control system effectors, as well as crew
intravehicular and extravehicular activities).
In addition to inertial loads, there are loads generated by forces directly applied
by cosmonauts (astronauts).
In addition to the above loads, the equipment installed on the outer surface of the
modules is subjected to the following loads:
jet plume impingement loads;
loads produced by a crew tether (when the tether is fastened to the equipment).
Inertial loads for the equipment are given in the form of operational linear and
angular accelerations of the center of gravity and about the center of gravity of the
equipment in the OXYZ coordinate systems, which are defined as follows:
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- SM: X axis of the OXYZ coordinate system is parallel to the XRS axis and is
pointed in the same direction. Y axis is perpendicular to the X axis, is parallel to
the YRS axis and is pointed in the same direction. Z axis completes the righthanded coordinate system;
- DC1 (located on the –YSM docking port): XDC1 axis of the ODC1XDC1YDC1ZDC1
coordinate system is parallel to the YRS axis and is pointed in the opposite
direction. YDC1 axis is perpendicular to the ХDC1 axis, is parallel to the ZRS axis and
is pointed in the opposite direction. ZDC1 axis completes the right-handed
coordinate system;
MRM2
(located
on
the
+YSM
docking
port):
XMRM2
axis
of
the
OMRM2XMRM2YMRM2ZMRM2 coordinate system is parallel to the YRS axis and is
pointed in the same direction. YMRM2 axis is perpendicular to the ХMRM2 axis, is
parallel to the ZRS axis and is pointed in the same direction. ZMRM2 axis completes
the right-handed coordinate system;
MLM: X axis of the OXYZ coordinate system is parallel to the YRS axis and is
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pointed in the same direction. Y axis is perpendicular to the X axis, is parallel to
the XRS axis and is pointed in the same direction. Z axis completes the righthanded coordinate system.
The following notation is used:
aX — acceleration in the X axis direction;
aYZ — acceleration in any transverse plane parallel to the OYZ plane;
εX — angular acceleration about X axis;
εYZ — angular acceleration about any transverse axis, which lies in a plane parallel
to the OYZ plane and goes through the X axis.
Recommended stiffness of fasteners for installed equipment shall correspond to
natural frequencies of no less than 20 Hz.
Operational linear and angular accelerations of the center of gravity and about
the center of gravity of the equipment located inside and outside ISS RS modules
with rigid attachment at interface points to the ISS RS in 6 degrees of freedom
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(that is, having the natural frequency of the attachment ≥20 Hz taking into account
the stiffness of the attachment point) are:
aX = 4.5 m/s²; aYZ = 6.0 m/ s²; εX = 0.2 rad/ s²; εYZ = 0.7 rad/ s².
For the equipment having a natural frequency of the attachment below 20 Hz,
operational values of linear and angular accelerations can be as high as the
following values:
aX = 22.0 m/s²; aYZ = 16.0 m/s²; εX = 1.3 rad/s²; εYZ = 7.0 rad/s².
When determining design loads, the safety factor shall be set equal to 2.
If need be, these accelerations can be updated for an individual piece of
equipment to take into account its specific attachment stiffness values, dynamic
characteristics of the primary structure of the equipment and its location on the ISS
RS structure.
The structure of the equipment installed inside ISS RS modules in the crew
transfer path, in addition to inertial loads (see paragraph 0.2) in accordance with
the requirements of approved document SSP 41163 shall withstand operational
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loads of 556 N (design value of 778 N) caused by the crew knocking directly
against the equipment.
The force shall be applicable in any direction to any spot on the equipment
accessible to the crew.
The structure of the equipment installed on the outer surface of the ISS RS in
close proximity to the crew transfer path, in addition to inertial loads (see
paragraph 0.2) in accordance with the requirements of approved document SSP
41163 shall withstand operational loads of 50 kgf (490.5 N) caused by the crew
inadvertently knocking against the equipment.
The force shall be applicable in any direction to any spot on the equipment
accessible to the crew.
In accordance with requirements of the approved document SSP41163, a
structure on the equipment which may serve as an attachment point for the hook of
a crew tether shall be capable of withstanding the design load of 200 kgf (1962 N)
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applied at the tether attachment point and acting in any direction accessible to the
crew.
In accordance with requirements of the approved document SSP 41163 the
structure of the equipment installed on the ISS RS outer surface (with the
exception of solar arrays and radiators), shall be capable of withstanding the
folowing local operational pressures generated by the thruster jets:
- normal pressure of 16.7 kgf/m²;
- circumferential pressure of 3.91 kgf/m².
The surface area to which these pressures are applied: up to 0.25 m2.
When equipment is installed on solar arrays and radiators, the values for
operational pressures are to be additionally defined.
Safety factor for operational pressures generated by thruster jets is to be set equal
to 2.
Operational number of the equipment loading cycles in the orbital phase of the
mission during docking, control system operation, as well as during intra- and
operational life is 15 years, is to be set in accordance with Table 6.2.5.
Table 6.2.5 Equipment cyclic loading spectrum
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extravehicular activities, for evaluating fatigue strength when the equipment
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Acceleration level, %
Operational number of acceleration
cycles
100—90
100
90—80
200
80—70
700
70—60
800
60—50
950
50—40
1 150
40—30
1 500
30—20
6,49E+6
20—15
1,40E+7
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Acceleration level, %
Operational number of acceleration
cycles
15—10
3,17E+7
10—2,5
8,05E+7
In this case, the 100-percent loading level is to be assumed to be equal to the
maximum loading level specified in paragraph 6.2.2.1.2.
As the equipment operational life increases, the number of loading cycles rises in
proportion to the increase in operational life.
6.2.3 Microgravity
6.2.3.1 Microgravity environment in RS modules
In accordance with Basic Requirements for ISS RS 27КСМ.0000-0ПЗ1, during
ISS microgravity flight modes, which, per specification SSP41163, shall be
maintained for 180 days per year in continuous 30-day intervals, RS modules
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provide the following microgravity environment:
- quasi-static accelerations (frequencies <0.01 Hz) of no more than 4 microg.
Quasi-static acceleration in excess of NASA requirements is due to a
disadvantageous position of the RS research modules with respect to the ISS center
of gravity;
- root-mean-square vibration acceleration levels in 1/3-octave frequency subranges
(averaged over time intervals of 100 s) do not exceed the levels shown in Fig.1.
Higher levels at frequencies ranging from 0.8 Hz to 5 Hz are due to the operation
of solar array actuators and US segment radiators (SARJ, TRRJ), crew exercise on
treadmill and stationary bicycle;
maximum amplitudes of transient accelerations caused by individual sources do
not exceed 10000 micro-g along each axis.
At frequencies of 10 Hz to 300 Hz requirements are met at distances of no less
than 1 meter away from the most powerful sources of disturbance, which include,
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in particular, compressors (for example, in air-conditioning systems), electric
pumps in the heating and cooling loops, toilet, control moment gyros.
Requirements may be updated for specific equipment locations.
6.2.3.2 Requirements for the equipment to maintain microgravity conditions
In order to comply with microgravity requirements for ISS RS and for ISS as a
whole during microgravity modes of the ISS flight (180 days per year in periods
stretching for 30 days each) every piece of equipment shall meet the following
requirements:
the equipment shall not produce quasi-stationary forces of more than 4 grams
acting for long periods of time (exceeding 30 seconds);
the equipment operation shall not cause accelerations with frequencies of up to 10
Hz:
in ISS RS modules, that are in excess of 10% of the levels shown in Fig. 6.2.3.1;
The operation of the equipment shall not cause accelerations with frequencies in
6.2.3.1 at the equipment attachment point;
the operation of the equipment shall not cause non-stationary accelerations with
amplitudes exceeding 1000 micro-g at the equipment attachment point.
6.2.3.3 Noise level limit requirements for the equipment
During its operation no piece of equipment shall generate noise in excess of 55 dB
at the distance of 1 m from the source.
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the range of 10 Hz to 300 Hz that are in excess of 10% of the levels shown in Fig.
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micro-g, 1/3-octave RMS value
Frequencies, Hz
Fig. 6.2.3.1
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6.2.4 Electrical requirements
6.2.4.1 Basic requirements
6.2.4.1.1 On-board power supply
Power is supplied to the equipment from a DC power source of the Electrical
Power Supply (EPS) system of an ISS module at nominal voltage of 28.5+/-0.5 V
across the EPS output terminals.
Steady-state supply voltage at the hardware input lies within the 23-29 V range.
The hardware shall operate normally when exposed to sudden changes in supply
voltage by +/-4.5V with a frequency no higher than 1 Hz and duration of no less
than 0.1 ms. In that case, the voltage at the hardware input shall remain within the
23-29 V range.
The on-board automatic equipment of the module provides bipolar switching of
the primary power supplied to the hardware, as well as protection of power lines
against possible overcurrent loads and short-circuits in the cabling and hardware.
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To assure an optimal selection of elements for power switching, overload
protection, and of power cabling parameters, the following inputs from the users
are needed:
the number of power feeders and their purpose;
timeline for applying and removing power in each feeder under nominal and offnominal situations;
characteristics of each power feeder, including:
the nature of the load (resistive, inductive, capacitive);
power consumption (watt, ampere) timelines for the equipment under all
operational modes, taking into account possible input voltage variations within
specified limits;
parameters of possible inrush currents in transient modes, including inrush current
amplitudes and their durations.
Acceptable values for inrush currents and their durations are additionally
specified in each individual case and are agreed during development of electrical
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schematics for the equipment interfaces with the module control system.
In general, inrush currents must not exceed the maximum steady-state operating
current multiplied by a factor of five. In that case the duration of the current pulses
must not exceed 20 ms when input power voltage varies within 23-29V.
Primary power circuits of the equipment shall have conductive coupling with
neither the case of the device, nor with the telemetry circuits.
The equipment must remain operational in case a primary power bus
inadvertently comes into contact with the device case.
Loads (devices) having individual protected power feeders must not have any
conductive coupling between themselves via the primary power plus buses (“+”).
As a rule, each device must be powered via a separate connector. Its type and pin
assignment are additionally defined during development of electrical schematic for
interfaces between the device and the onboard automatic equipment.
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6.2.4.1.2 Electrical insulation resistance.
Resistance of electrical insulation between primary power supply circuits (both
positive and negative) and equipment case shall be:
no less than 20 MOhm at relative humidity of (45…80) percent and ambient
temperature of (+15…+35) deg.C.
no less than 1 MOhm at relative humidity of 95+/-3 percent and ambient
temperature of 20+/-5 deg.C.
Insulation resistance measurements are taken at 30V DC.
6.2.4.1.3. Electric strength of the insulation.
Resistance of insulation between primary power circuits (both positive and
negative) and the case of the equipment, as well as between any electrically
isolated circuits, shall comply with requirements for electrical insulation resistance
after application to these circuits of a 200V test voltage (effective or constant
value).
The test voltage shall not exceed allowable test voltages for components used in
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the equipment.
6.2.4.1.4 Requirements for data interfaces with the ISS RS systems
The equipment, which has data interfaces with the ISS RS systems, shall be safe
for the Russian hardware and control systems that are sharing the interfaces.
The equipment to be used onboard RS, shall be classified by the presence of the
following interfaces:
- electrophysical interfaces, which determine parameters of interfacing with
electrical circuits for data transmissions. Such an interface must be subjected to the
following operations:
device connectors shall be mated to the actual on-board cabling to check their
wiring;
electrical circuit parameters shall be measured in the actual onboard cabling and
evaluated to make sure they lie within tolerances for data transmission lines (MIL
STD1553B, ETHERNET, etc.)
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performance tests shall be run via connected interfaces per user’s manual for the
device.
- information and logic interfaces determining protocols for data transmissions in
the network to which the new user is connected, and clearly defining its functions.
To be checked are:
logic functions of the device as a user of a data network (controller or slave
device);
the impossibility for the user to corrupt formats and data content in an interface
unrelated to the user;
correctness of the user response to the set of commands contained in the interface.
- software interfaces between newly-added equipment and the user providing
protection of RS systems against off-nominal situation.
To be checked in such an interface (belonging, for example, to a laptop computer)
are hardware protection against false and invalid commands and operator’s
instructions.
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The equipment shall be checked in real onboard environment in order to obtain
actual measurement results from the flight model of the onboard cabling against
the background of active operation mode of the onboard systems over applicable
interfaces.
In verifying compliance with data security requirements, a special emphasis
should be placed on standard interchangeable components and storage medium
containing installation or autorun software (cartridges, hard drives, floppy disks,
etc.)
The use of such elements onboard shall be tightly controlled and measures shall
be in place to guard against their unauthorized use in the RS equipment.
To be checked is whether the protection is in place against:
unauthorized launch of software and its installation in RS computers;
corruption of basic software of the RS equipment supporting control system
operation.
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Protective/warning labeling must be made and applied to interchangeable
components to limit the list of devices where these could be installed.
6.2.4.2. Electromagnetic compatibility
6.2.4.2.1 Electromagnetic interference generated by scientific equipment
6.2.4.2.1.1 Low-frequency noise
Peak voltage values for low-frequency noise (Upeak), generated by equipment in
the ±28V power supply circuits of the ISS RS modules shall not exceed the values
given in Fig. 6.2.4.2-1.
f [kHz]
Fig. 6.2.4.2-1 – Low-frequency noise
Where: f is the frequency (kHz)
Measurement band shall be at least:
10 Hz in the range of 30 Hz to 1 kHz;
100 Hz in the range of 1 to 10 kHz.
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Upeak[dBμV]
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6.2.4.2.1.2 RF noise
Peak voltage values for RF noise shall not exceed values given in Fig.6.2.4.2-2.
Upeak[dBμV]
f [MHz]
Fig. 6.2.4.2-2 – RF noise
Measurement band shall be at least:
1 kHz in the range of 0.01 to 0.15 MHz.
10 kHz in the range of 0.15 to 30.0 MHz.
6.2.4.2.1.3 Electric field strength of radiated RF noise.
Equipment installed onboard ISS RS shall not generate noise exceeding the
limits specified in Fig. 6.2.4.2-3. Above 30 MHz, compliance with the limits shall
be assured for both horizontally and vertically polarized waves.
Epeak[dBμV/m]
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100 kHz in the range of 30 to 100 MHz.
f [MHz]
Fig. 6.2.4.2-3 – Electric field strength
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Measurement band shall be at least:
1 kHz in the range of 0.01 to 0.15 MHz;
10 kHz in the range of 0.15 to 30.0 MHz;
100 kHz in the range of 30 to 1000 MHz;
1 MHz above 1000 MHz.
These requirements do not apply to RF noise generated by emissions from radio
transmitter outputs.
6.2.4.2.2. Immunity of equipment to electromagnetic noise
6.2.4.2.2.1 Low-frequency noise
Peak voltage values for low-frequency noise in the ±28V power supply circuits
of the module are given in Fig. 6.2.4.2-4.
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Upeak[V]
f [MHz]
Fig. 6.2.4.2-4 – Low-frequency noise
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6.2.4.2.2.2 Industrial RF noise
Peak voltage values for industrial RF noise in the ±28V power supply circuits of
the ISS RS modules are given in Fig. 6.2.4.2-5.
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Upeak[dBμV]
f [MHz]
Fig. 6.2.4.2-5 – Industrial RF noise
6.2.4.2.2.3 Pulse noise
The effect of the pulse noise is characterized by the following parameters:
- amplitude and duration of a test pulse are given in Tables 6.2.4.2-1 and 6.2.4.2-2;
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- pulse edge durations are no more than 5% of the pulse duration;
- pulse rate is 1 Hz for a period of 1 minute (or for a period of time needed to
evaluate performance).
Table 6.2.4.2-1 Characteristics of pulse noise between power buses
Parameter
Pulse
Value
duration,
microseconds
Pulse amplitude, V
100
300
500
+ 15
+ 15
+ 10
+ 10
– 15
– 15
– 10
– 10
Table 6.2.4.2-2 Characteristics of pulse noise between each of the power buses and
the case
Value
Parameter
Pulse
duration,
microseconds
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50
50
100
200
300
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Pulse amplitude, V
+ 35
+ 35
+ 10
+ 10
– 35
– 35
– 10
– 10
Peak values of electric field strength inside ISS RS modules are given in Fig.
6.2.4.2-6.
Epeak[dBμV/m]
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f [MHz]
Fig. 6.2.4.2-6 – Electric field strength inside the module
6.2.4.2.2.4 Radiated RF noise outside ISS
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Peak values of electric field strength outside ISS RS are given in Fig. 6.2.4.2-7.
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Epeak[dBμV/m]
f [MHz]
Fig. 6.2.4.2-7 – Electric field strength outside ISS RS
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Field strength may vary from 134 dBmicroV/m to the specified value depending
on the device location.
The radiated signal is modulated in amplitude with 1 kHz frequency and 50%
depth of modulation.
6.2.4.2.3. Requirements for radio receivers, radio transmitters and AFD
Specifications of radio receivers, radio transmitters and antennas meet the
requirements of Radio Regulations and recommendations of International Radio
Consultative Committee (IRCC).
Specifications of radio receivers, radio transmitters and antennas are checked for
compliance with the standards of Radio Regulations using procedures, which meet
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the requirements of IRCC Recommendations.
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6.2.5. Labeling
6.2.5.1. Special labeling for ISS
In addition to the labeling usually used by the equipment developer or vendor,
alphanumeric and barcode labels must be applied. Equipment delivered in a
package shall have the proper labeling on the package.
6.2.5.1.1 Alphanumeric labeling.
Alphanumeric labeling in two languages, the types and methods of labeling,
specifics of using Rusian and English are described in the joint standard SSP
50094 (Section 4.6.7).
6.2.5.1.2 Barcoding.
In order to set up inventory accounting of every piece of equipment and cargo,
both delivered to ISS RS, and returned/disposed of from the space station, the ISS
Cargo Traffic and Inventory is used. Each item included into block-by-block list
which is entered into and then stored in the information system ISS Cargo Traffic
and Inventory. To automate the inventory keeping processes, barcoding is used;
therefore, all the items intended for use onboard ISS shall have barcode labels of
one of the sizes with different information content:
- barcode only, also duplicated in numerical format,
- name and barcode, duplicated in numerical format,
- system designation, name and barcode, duplicated in numerical format.
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and into cargo traffic shall have a description following a pre-defined data set,
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6.2.6. Ionizing radiation
6.2.6.1 Integral energy spectra of electrons and protons in Earth radiation belts at
solar maximum are given in Table 6.2.6-1.
Table 6.2.6-1
Proton energy,
Proton stream,
MeV
cm-2⋅day-1
MeV
cm-2⋅day-1
0,04
1,5E+10
1,00
3,9E+06
0,10
9,1E+09
2,00
2,3E+06
0,30
1,5E+09
3,00
2,0E+06
0,50
4,4E+08
4,50
1,7E+06
1,00
1,2E+08
6,00
1,6E+06
1,50
5,7E+07
10,00
1,4E+06
2,00
2,6E+07
20,00
1,2E+06
3,00
4,8E+06
100,00
5,9E+05
4,00
6,9E+05
300,00
6,9E+04
5,00
6,9E+04
500,00
8,2E+03
5,40
2,2E+04
700,00
3,7E+02
Note – These data are used to evaluate integrity of materials and elements located
outside ISS RS, for which radiation hardness parameters are specified in the form of
spectral characteristics
6.2.6.2 Absorbed doses of space electrons and protons for materials and equipment
located inside ISS RS per 1 year of flight tests for various values of mass thickness
of a spherical shielding under exposure to omnidirectional irradiation (solid angle
4π) are listed in Table 6.2.6-2.
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Electron energy, Electron stream,
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Table 6.2.6-2
Absorbed dose, rad/year
Mass
thickness, g/cm2 electrons
protons
total
3,0E-01
2,4E+03
9,3E+01
2,5E+03
5,0E-01
1,1E+03
7,5E+01
1,2E+03
7,0E-01
5,9E+02
6,9E+01
6,5E+02
1,0E+00
2,4E+02
6,2E+01
3,0E+02
2,0E+00
9,0E+00
4,9E+01
5,8E+01
3,0E+00
1,6E+00
4,3E+01
4,4E+01
5,0E+00
1,0E+00
3,4E+01
3,5E+01
1,0E+01
6,2E-01
2,3E+01
2,4E+01
6.2.6.3 Absorbed doses of space electrons and protons for materials and equipment
installed outside ISS RS per 1 year of flight tests for various values of mass
(solid angle 2π) are listed in Table 6.2.6-3.
Table 6.2.6-3
Absorbed dose, rad/year
protons
total
1,0E-05
4,0E+05
3,1E+04
4,3E+05
1,0E-02
7,3E+04
2,3E+02
7,4E+04
5,0E-02
9,3E+03
7,2E+01
9,4E+03
7,0E-02
5,2E+03
6,1E+01
5,3E+03
1,0E-01
2,8E+03
5,2E+01
2,8E+03
2,0E-01
8,7E+02
3,9E+01
9,1E+02
5,0E-01
1,8E+02
2,8E+01
2,0E+02
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thickness, g/cm2 electrons
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Mass
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thickness of a spherical shielding under exposure to unidirectional irradiation
6.2.6.4 All radiation doses shall be recalculated for the specified duration of flight
tests of instruments and equipment assuming that they are proportional to the
duration of flight tests. Dosage safety factor (with respect to the listed data or data
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obtained by recalculation for other operation life) for the developed (purchased)
equipment shall be no less than 2.
6.2.6.5 In case the hardware is reused guaranteed in-orbit operation time is
determined by taking into account the total planned duration of flight tests.
6.2.6.6 Ionizing radiation responsible for soft and hard failures in the ISS RS
equipment is characterized by integral spectra of linear energy transfer for heavy
charged particles and differential energy spectra for protons.
Table 6.2.6-4 provides integral spectra of linear energy transfer for heavy charged
particles.
Table 6.2.6-4
Linear energy
Linear energy transfer for the heavy charged particles source
Galactic cosmic rays at solar
Solar cosmic
Solar cosmic
MeV⋅cm2/mg
minimum (flux averaged over an
rays (total
rays (peak
orbit), particles/(cm2⋅days)
flux),
flux),
particles/cm2
particles/(cm2⋅d
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transfer
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7,1E+01
6,7E+04
4,2E+06
7,0E-01
5,5E+01
4,9E+04
2,4E+06
1,0E+00
4,0E+01
3,4E+04
1,3E+06
3,0E+00
1,2E+00
8,4E+03
1,9E+05
5,0E+00
5,1E-01
4,1E+03
7,5E+04
1,0E+01
1,4E-01
1,1E+03
1,5E+04
2,0E+01
2,7E-02
2,0E+02
2,1E+03
3,0E+01
1,7E-05
1,8E+00
1,4E+01
4,0E+01
9,8E-07
1,5E-01
1,7E-01
5,0E+01
4,1E-07
7,7E-02
7,7E-02
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5,0E-01
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Linear energy
Linear energy transfer for the heavy charged particles source
transfer
Galactic cosmic rays at solar
Solar cosmic
Solar cosmic
MeV⋅cm2/mg
minimum (flux averaged over an
rays (total
rays (peak
orbit), particles/(cm2⋅days)
flux),
flux),
particles/cm2
particles/(cm2⋅d
ays)
7,0E+01
7,5E-08
1,6E-02
1,6E-02
9,0E+01
6,0E-09
1,4E-03
1,4E-03
Table 6.2.6-5 provides differential energy spectra for protons.
Table 6.2.6-5
Earth radiation belts at
Earth radiation belts at
Solar cosmic rays
solar minimum (flux
solar minimum (peak
(peak flux),
averaged over an orbit),
flux at SAA),
protons/(cm2⋅day⋅M
protons/(cm2⋅day⋅MeV)
protons/(cm2⋅days⋅MeV) eV)
30
2,0E+04
8,2E+05
3,5E+08
50
1,7E+04
1,1E+06
1,4E+08
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70
1,4E+04
1,1E+06
6,6E+07
100
1,1E+04
8,4E+05
2,7E+07
140
6,9E+03
5,5E+05
1,1E+07
Взам. инв. №
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200
3,2E+03
2,6E+05
3,3E+06
400
3,9E+02
3,5E+04
2,6E+05
600
6,4E+01
4,8E+03
4,4E+04
Energy,
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MeV
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6.2.7. Requirements for materials
The below requirements apply to all materials listed in the full list of nonmetallic
materials. The full list of nonmetallic materials, which are in contact with the
pressurized cabin atmosphere, includes:
- structural materials used in fabrication of casings;
- fasteners, circuit boards, covers, sheathes, etc.;
- varnishes, paints, compounds, adhesives;
- insulating materials for wires, cables, harnesses.
The list has the official document status and must be signed by a responsible
representative of the international partner.
6.2.7.1. Materials fire safety
Materials shall not be capable of spontaneous ignition.
Materials shall not be capable of maintaining combustion either on their own or
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when integrated into another product or system.
Fire-prevention measures
- all the circuit boards shall be covered on both sides with silicon or another
approved material, if they are not completely enclosed in a metal container and if
they are not otherwise accepted by the Russian side;
- all transformers or coils are fully enclosed in a metallic material or an approved
potting compound. Exceptions from this rule must be specifically approved by the
Russian side;
- all the wires are covered with Teflon, all wires and cables are protected or routed
in such a way as to rule out any possibility of wire damage which may cause
sparks or short-circuits;
- circuit protection shall be provided in order to avoid possible circuit overheating
in off-nominal situations (for example, short circuit, as a result of an abrupt change
or a delay);
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- insulation of power cables shall be heat resistant; connectors that do not have
such protection shall be shielded with housings made of nonflammable materials.
6.2.7.2. Materials toxicity
It is forbidden to use for the hardware structure any toxic or hazardous materials
which may release gases detrimental to the crew in concentrations dangerous to the
crew.
Materials of the equipment located on the outer surface of ISS shall meet the
standards for outgassing in vacuum with respect to loss of mass (no more than 1%)
and the amount of condensing substances (no more than 0.1%), determined per
GOST Р50109-92 or another standard similar to ASTME595.
6.2.7.3 Microbiological stability of the materials
Equipment destruction to a limit posing a threat to health and life is not allowed.
Microbial resistance must be specified for each listed material.
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Equipment and its container must be disinfected at the launch site, unless
otherwise specified by the Russian side.
Usual methods of hardware disinfection include the use of liquid hydrogen
peroxide and radiation. This usual processing applies to all external surfaces. This
method can be revised if there is any danger of damaging the hardware.
Disinfectant resistance shall be specified for each listed material.
6.2.7.4 Aging of materials
An estimate of the equipment susceptibility to ageing shall be made by means of
non-metallic materials analysis. Assessed on a regular basis shall be devices
intended for multiple use in flight, as well as items and materials, which are used
as samples as defined by the Russian side. Periods of time during which the
properties of the materials remain preserved shall be specified for each listed
material.
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6.2.8 Protection against static electricity
6.2.8.1 Protection of electronic equipment against static electricity during its
installation and hookup to onboard cables shall be provided by taking the
following measures:
- bonding the equipment, that is, providing a secure electrical connection between
the cases of the devices and the body of the ISS RS modules;
- removing static electricity from the operator during electronic equipment
installation using an antistatic wristband connected to grounding points available in
the ISS RS modules.
- removing static electricity from onboard cables during their connection to
electronic equipment by using a multipurpose device 17КС.300Ю9052-0.
6.2.8.2 Hardware bonding is provided by means of bonding straps, equipped with
which must be casings of the devices to be installed inside pressurized
compartments.
and to the body of the pressurized compartment by means of a fastener (a screw),
or by mating disconnectable bonding straps.
It is acceptable to bond devices via an electrical interface with the ISS RS
onboard systems, where one of the electrical circuits which has galvanic coupling
with the pressurized compartment body is assigned to serve as a bonding wire.
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The bonding is established by connecting a bonding strap to the device casing
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6.3 Safety requirements
6.3.1 General requirements
6.3.1.1 General requirements are applicable to SE design and development and are
aimed at assuring the required level of SE safety both for the crew and for other
hardware on the vehicle.
6.3.1.2 During its operation and storage the SE shall not crate dangerous situations
or the risk of such situations occurring shall be minimized.
The dangerous situations, as applied to SE, are classified according to the hazard
level as follows:
- Critical Hazard. The hazard which, if it occurs, may result in such damage to the
vehicle equipment as does not preclude its further operation, or in a nondisabling
injury to the crew, or in the need to use unplanned procedures, which affect the
operation of onboard systems of the Russian spacecraft;
- Catastrophic Hazard. The hazard which, if it occurs, may result in the loss of the
6.3.1.3 Development of SE and its utilization operations shall meet the following
6.3.1.3.1 Tolerance to failures resulting in dangerous situations
operator error can result in a catastrophic hazard;
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requirements in order to minimize the risk of occurrence of hazardous situations.
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vehicle or in a disabling or fatal injury to the crew.
SE shall be designed in such a way that:
- no combination of two failures or two operator errors, or one failure and one
- no single failure or single operator error can result in a critical hazard.
6.3.1.3.2 Design for Minimum Risk
This requirement consists in assuring/selecting the necessary safety factors,
scope of developmental and verification tests to assure high strength properties,
survivability, resistance to various kinds of loads. Safety factors and scopes of
testing are determined by standards depending on the types of equipment and
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specifics of its use. Typical examples of the use of the design for minimum risk are
selection of safety factors, of safety margins when analyzing the structural strength
of the body of sealed pressurized vessels, of load-bearing structural elements.
6.3.1.3.3 Safety taking into account vehicle/ISS services
6.3.1.3.3 Safety without taking into account vehicle/ISS services
SE shall be capable of taking care of its safety by itself without any servicing on
the part of the vehicle/ISS even in case of a contingency onboard the vehicle/ISS.
6.3.1.3.3.2 Safe With Vehicle Services
If requirements of safety without servicing (paragraphs 6.3.1.3.3.1 or 6.3.1.3.3.2)
are difficult to implement for technical reasons, the requirement of safety with
servicing on the part of the vehicle must be met.
In this case the overall safety of the SE together with the vehicle equipment that
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is being used shall meet the requirements of paragraph 6.3.1.3.1 or 6.3.1.3.2 of this
document with respect to hazardous situations.
6.3.1.3.4 Hazardous functions control
6.3.1.3.4.1 A function, which, if inadvertently triggered, may result in a critical
hazard, shall have two independent inhibits for the entire time while the likelihood
of this hazard exists. In those cases where the loss of function may result in a
critical hazard no single failure shall result in the loss of this function.
6.3.1.3.4.2 A function, which, if inadvertently triggered, may result in a
catastrophic hazard, shall have at least three independent inhibits for the entire time
while the likelihood of this hazard exists. There shall be a capability to monitor the
status of at least two of the three required inhibits . In those cases where the loss of
function may result in a catastrophic hazard no two failures shall result in the loss
of this function.
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6.3.1.3.5 Caution and warning devices
In those cases where the occurrence of a hazardous situation requires urgent
action on the part of the crew (fire, depressurization, toxic release, etc.), devices
shall be used for providing a timely warning to the crew about these hazardous
situations and/or for switching off the SE (system hardware).
6.3.1.3.6 Special procedures
Procedures supporting the SE safety shall contain: Safety instructions for
equipment handling, procedures to be followed in case of possible SE failures.
6.3.2 Design requirements
6.3.2.1 Materials
Materials used in SE structure, including packaging, shall meet the following
safety requirements.
a) Toxicity
The materials in use that are in contact with the spacecraft environment shall not
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release into the environment any toxic and noxious chemical substances in
concentrations constituting a threat to the life of the crew and spacecraft operation.
Care must be taken to avoid the use in the SE hardware of any chemical
substances which do not come into direct contact with the environment, but may
pose toxicity threat if they leak into the spacecraft atmosphere (for example, may
cause crew skin or eye irritation or affect spacecraft equipment). If, nevertheless,
they need to be used, requirements of paragraph 6.3.1.3 shall be met in accordance
with the hazard level.
b) Fire safety
Materials used under normal operating conditions and in case of cabin
depressurization:
shall not become a source of fire or explosion;
shall be heat-resistant and shall not sustain combustion neither by themselves, nor
when used within hardware assemblies and systems;
shall not release flammable gases and form combustible or explosive mixtures.
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c) Microbiological safety
The materials in use shall not be a source of microbiological contamination of
the spacecraft atmosphere and structural elements, and shall not create a propitious
environment for microorganism development, and shall be resistant to
biodegradation.
d) Operation in vacuum
SE located outside the pressurized cabin of the spacecraft shall be resistant to
exposure to heat and vacuum and shall not be a source of contamination of the
space around and inside the module under normal environmental conditions.
Non-metallic materials, as well as coatings used for sealing and insulating
electrical circuits shall be tested for mass loss and volatile matter content under
exposure to heat and vacuum and have mass loss of no more than 1% and volatile
matter content of no more than 0.1%.
Thermooptical properties of the surfaces of the SE located outside the
pressurized cabin of the vehicle shall not affect the vehicle safety: exceed the
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allowable thermal conditions, as well as create glare or flood exposure of lightsensitive elements of the vehicle.
6.3.2.2 Structure
6.3.2.2.1 SE structure shall be designed taking into account actual loads and safety
factors accepted for the vehicle.
6.3.2.2.2 Materials used for the SE structural elements, including fasteners, shall be
reviewed for corrosion cracking under stress (fatigue strength). As far as possible,
alloys with high resistance against stress corrosion cracking shall be used.
6.3.2.2.3 SE structure shall be designed for maximum differential pressure
occurring during ascent, descent or activities involving emergency depressurization
and subsequent re-pressurization of the vehicle/ISS. The vents in the vented
compartments shall have such surface areas as to assure structural integrity while
providing the maximum rate of change for the pressure.
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6.3.2.2.4 Open surfaces on SE shall be smooth and have no burrs.
6.3.2.2.5 The structure of the SE and the packing used in flight shall have no
breakable materials or shall preclude fragments of such materials spreading beyond
casings or packing of the SE hardware in case of failures both under normal
environmental conditions and in case of cabin depressurization.
6.3.2.3 Electrical safety
SE electrical circuits shall be protected against possible overloads and shortcircuits (and potential ignition) by correct choice of wire gauge, wire insulation,
circuit breakers, etc., etc.
There shall be no electrical contact between the vehicle body and electrical
circuits of the SE.
SE hardware shall be designed in such a way as to protect the crew against an
inadvertent contact with electrical circuits. In case the SE uses voltages above 30
catastrophic hazard level.
Leak current during contact with the SE being serviced shall not exceed 0.07 mA
DC for a normal contact with a human being.
shall have a voltage difference exceeding 40 mV at frequencies of 1000 Hz or
6.3.2.4 Radiation safety
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No two electrically conductive surfaces within the reach of a member of the crew
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V, measures shall be taken to protect the crew against electric shock per
lower, measured across the resistance of 1000 Ohm.
All the SE elements which produce ionizing radiation shall be identified and their
use shall be the subject of a special discussion and agreement with RSC Energia.
6.3.2.5 Living space safety
SE shall not be the source of elements contaminating internal compartments of
the vehicle. To assure the cleanliness of the living space inside the compartments,
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liquid, loose and paste-like substances shall be contained within a reliable sealed
package.
The SE structure shall guarantee that the amount of outgassing will not exceed
maximum allowable concentrations, that is, the gas composition will not be
disturbed. SE shall not be a source of strong smell.
The acceptable noise level during SE operation shall not exceed 65 dBA at the
distance of 1 m.
6.3.2.6 Non-ionizing radiation
The use of SE containing sources of laser radiation, as well as containing
permanent magnets, shall be agreed with RSC Energia.
6.3.2.7 In-flight handling safety
6.3.2.7.1 SE shall have interlocks to prevent inadvertent operation or a change in
configuration caused by crew actions during handling. For example, locking of
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switches can be achieved by means of protective covers. If need be, onboard
documentation (instructions) for the crew shall be prepared.
6.3.2.7.2 SE external structure and packaging shall have no sharp or protruding
elements.
6.3.2.7.3 Touch temperature
Surfaces having temperature of 40 to 45оC, with which the crew may come into
an inadvertent contact, shall have appropriate caution plates.
Surfaces with temperatures above 45оC shall have protective devices or
insulation to prevent crew contact.
Surfaces with temperatures below 5оC, with which the crew may come into
contact shall have protective devices or insulation to prevent such contact, as well
as caution plates.
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6.3.2.7.4 Spherical radii of outer corners and edges
The SE hardware shall have rounded corners and edges, or means of protecting
the corners in accordance with Table 6.3.2.7.4
Table 6.3.2.7.4 Spherical radii of outer corners, planes and edges
Structural
Radius, mm
elements/fragments
Outer
Inner
Holes, panels, covers
6,4
3,0
Preferable
(corner radii in the panel
3,0
1,5
Minimal
Note
plane)
Open corners
13,0
Minimal
Open edges:
(1)
thickness of 2.0
≥1,0
(2)
(3)
and more
full radius;
thickness of
Smoothed
0.5 to 2.0 mm
out or
thickness of less
rounded off
Minimal
Small pieces of hardware 1,0
Note: Instead of a spherical radius, a chamfer 450 by 1.5 mm (minimum) with
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than 0.5 mm
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mm
Minimal
manipulated by hand
wearing a spacesuit glove
smoothly beveled edges. The chamfer width shall be selected in accordance with
approximate values for spherical radii described above.
6.3.2.7.5 Screws and bolts
Bolts and screws with threaded parts may protrude by 2-3 mm but by no more
than 1 turn of the thread, and have means of protection, which do not interfere with
installation or removal of the bolt or the screw. Fasteners shall be captive.
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6.3.2.7.6 Moving parts
Moving parts (fans, belt-drives, turbine wheels, etc.), which may cause crew injury
or damage to SE through inadvertent contact or capture, shall have guard rails or
other protective devices.
6.3.2.7.7 Pitch or capture points
Controls, locks and other similar devices shall be designed and positioned in
such a manner as to prevent them from pinching parts of the crew spacesuit. If
need be, protective covers or enclosures shall be used.
SE hardware located outside habitable compartments, which during its operation
may produce a gap of less than 3.5 mm between the hardware and adjacent
structures shall be designed in such a manner as to prevent pinching or capture of
parts of spacesuit of a crew member performing an EVA.
6.3.2.7.8 SE installation
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Internal and external SE located along the crew translation paths and installed in
the working areas shall be able to withstand loads caused by crew activities,
including EVA.
6.3.2.8 Chemical power sources (batteries and storage batteries, electrochemical
cells)
The structure, materials and electrolytes used in the chemical power sources shall
meet requirements of paragraphs 6.3.1 and 6.3.2.1 of these requirements.
Chemical power sources and/or devices, which use them, shall have protection
against charging/discharging overcurrents, excessive temperature and devices
preventing discharges of electrolyte into the atmosphere of the vehicle.
The use of chemical power sources with more than 120 watt-hour capacity or
having especially toxic electrolytes
(such as theonylchloride lithium) is only
allowed upon agreement with RSC Energia
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Fig. 6.4 – A typical ISS RS crew day plan
6.4 A typical crew day plan during implementation of the standard mode
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6.5 Acoustic environment in the ISS RS modules
ISS modules are filled with noisy equipment, which operates continuously or for
long periods of time throughout the day and constitutes the backbone of the life
support systems. Such systems include: ventilation system, air conditioning
system, system for filtering harmful contaminants from the air, thermal control
system. Immediate sources of continuous noise are such elements of these systems
as fans, pumps, compressors, electrical drives for various purposes, mechanically
operated controllers, etc.
There are no international standards for noise levels inside habitable compartments
of spacecraft. In Russia there is a currently operating standard GOST R 50804-95
Crew Habitable Environment in a Manned Spacecraft. General Medical and
Technical Requirements recommendations of which, as applicable to the
International Space Station became a constituent part of Russian-US standard Joint
NASA/RSA Document On Specifications And Standards For The ISS Russian
Segment (SSP 50094) which was put in force in 1996.
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Used as the unit for evaluating human exposure to noise usually is the level LA
measured in decibels with scale A correction (dBA) which corrects for nonuniformity in an average human’s hearing acuity for sounds with the same level at
different frequencies within the audible range. It should be noted that the noise
level LA does not always give a precise characterization of the noise perception and
only serves as its approximate criterion.
For missions lasting more than 30 days, SSP 50094 establishes maximum
allowable acoustical levels in habitable compartments of the spacecraft and
methods of their measurement. According to SSP 50094, noise from continuously
operating sources is restricted to the limiting acoustic pressure spectrum (in dB) in
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octave frequency bands with their geometric average values ranging from 31.5 to
8000 Hz, as well as to integral sound levels (LA ) and equivalent sound levels
(LA,equ.), measured on the A scale in dBA. By equivalent noise level (LA,equ.) is
meant a value that is equal to the constant noise level (LA), which has the same
energy. With continuously operating noise sources, the acceptable noise pressure
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levels in octave frequency bands, acoustic levels and equivalent acoustic levels in
habitable compartments of the Russian Segment of the International Space Station
established separately for the crew working and sleeping modes are listed in Table
6.5.1.
With the joint operation of the primary (continuous) and additional
(intermittent) noise sources, the aim is to make sure that the equivalent noise
level for the period of active crew operations (16 hours) does not exceed 60 dBA.
And in all the cases where the equivalent sound level exceeds 60 dBA taking into
account additional intermittent noise sources, the noise from all these sources shall
be limited by operating time parameters of the additional equipment per Table
6.5.2.
Table 6.5.1
Acceptable sound pressure levels in octave frequency bands and equivalent
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scale-A corrected sound levels for continuous noise in the habitable compartments
of manned spacecraft
Mission duration of
Sound pressure levels (dB) in octave
Equivalent
more than 30 days
frequency bands
noise levels
and sound
levels
expressed in
dBA
Center
frequencies 63
(Hz)
12
25
50
5
0
0
1k
2k
4k
8k
Work (16 hours)
79
70
63
58
55
52
50
49
60*
Sleep (8 hours)
71
61
54
49
45
42
40
38
50*
*) These noise levels are given as rounded values, calculated for octave frequency
bands.
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Taking into account additional intermittent noise sources, the total level of sound
per day measured in dBA, and duration of exposure to it per day shall not exceed
acceptable levels given in Table 2, with the equivalent level for the crew active
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period not exceeding 60 dBA.
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Table 6.5.2
Maximum acceptable sound levels
over a 16-hour working period in habitable compartments and in locations where
the crew stays for short periods of time during operation of additional noise
sources, depending on the time of exposure to them,
Maximum exposure time (hours)
Total acoustic pressure level (dBA)
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over a 16-hour working period
4
63
2
66
1
69
0.5
72
Acoustic pressure levels in octave frequency bands from additional noisy
equipment delivered onboard the manned spacecraft (not covered by
requirements of Tables 1 and 2, but included in the payloads/scientific
duration of their activation. In order to evaluate the effect of the additional
equipment, in each specific case an acoustic analysis is conducted of the
of noise.
levels allowable per sanitary and hygienic standards during crew active duty and
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mission phase, which includes this source and all the other significant sources
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equipment) shall be blow those given in Tables 1 and 2 depending on the
The noise exposure through an entire multi-day mission, which exceeds the
even during sleep, along with other adverse spaceflight factors results in higher
crew fatigue and loss of hearing acuity. This circumstance, and increasingly
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difficult recognition of voice commands and audio warning signals against the
noise background also affect flight safety.
6.5.1 Acoustic environment in the SM
SM is designed in such a way as to allow the crew to continuously reside in it,
and, in particular, this module has crew quarters (cabins) where crew members can
sleep. After a number of measures have been implemented to reduce levels of
noise from these noise sources, which included installing various devices (noise
suppressors, vibration insulators, sound absorbent mats, vibration damping
coatings), at present noise levels in the SM working compartment generated by the
continuously operating onboard equipment, according to in-flight acoustic
measurement results, are:
- in the central station area ― 65 … 67 dBA;
- in the wider part of the working compartment ― 64 … 65 dBA;
- in crew cabins (with fully closed doors) ― 52 … 55 dBA.
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It should be kept in mind that the controlled noise levels shall correspond to a
mode where only continuously operating onboard equipment is working in normal
operational mode in the absence of noises from intermittent sources and in the
absence of any noise generated by crew members. (It should be noted that the noise
generated by crew activities is not controlled in any way at all.)
6.5.2 Acoustic environment in DC1
DC1 design allows the crew to stay in it for short periods of time.
Based on the results of ground pre-flight acoustic tests, the noise level in the
central part of DC1 was ~70 dBA. There were no plans to implement measures to
reduce levels inside DC1.
In-flight acoustic measurements produced lower values:
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66 … 68 dBA, which may be due to a change in acoustic properties of the internal
volume of the module resulting from the fact that it is partially cluttered up with
various objects (especially, soft objects). During ground tests the module was
empty.
6.5.3 Acoustic environment in MRM2
MRM2 design allows the crew to stay in it for short periods of time.
MRM2 geometry and configuration of its continuously operating onboard
equipment are similar to DC1.
Based on the results of ground pre-flight acoustic tests, the noise level in the
central part of MRM2 was ~70 dBA.
6.5.2 Acoustic environment in MRM1
MRM1 design allows the crew to stay in it for short periods of time.
central part of MRM1 was, when averaged along the length of the working
compartment, ~72 dBA.
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Based on the results of ground pre-flight acoustic tests, the noise level in the
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6.6 Cargo integration
6.6.1 General
Progress and Soyuz spacecraft deliver to the space station “dry” cargoes in
pressurized compartments. In addition to this, Progress can deliver to the station
liquids and gases. If need be, cargoes can be placed outside the pressurized
compartment.
Cargoes transported by the vehicles shall meet requirements set fourth in this
document.
Documentation shall be drawn up for all the cargoes (including the packing used
in flight) to state that the cargoes are allowed to be stored and operated onboard the
International Space Station and transported onboard the spacecraft, as well as to
certify the flight safety of the cargo. The cargo accompanying documentation and
technical manuals must be approved by RSC Energia.
In order for RSC Energia to perform work on integration of the cargo and to
study the feasibility of its accommodation onboard the spacecraft, a documentation
6.6.1.1 For all cargoes:
- envelope and installation drawing;
- transportation safety data package;
- accompanying documentation.
6.6.1.2 For active cargoes, additional documentation is provided in accordance
with agreements between RSC Energia and the cargo owner on its integration and
delivery, including:
- electrical interface, including electrical consumption;
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package is to be submitted, which contains:
- heat release
- cargo operations timeline;
- special requirements (activation, telemetry, etc.)
Requirements for envelope and installation drawing are given in paragraph 6.7.1.
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Requirements for transportation safety data package are given in paragraph 6.7.2.
Requirements for accompanying documentation are given in paragraph 6.7.3.
The cargo owner is responsible for validity of provided data and for cargo
compliance with the specified requirements.
Cargo delivery, removal and return are only performed in case of positive results
of RSC Energia’s studies of cargo compliance with the requirements and feasibility
of their accommodation and fastening. The studies are conducted on the basis of
the documentation provided by the cargo owner/curator.
6.6.2 Project Management
Cargo integration project management consists of the following steps:
1)
reviewing and approving the documentation package on the cargoes
manifested for delivery on a specific spacecraft;
active cargoes require matching the interfaces;
2)
engineering analysis − development of the configuration for cargo
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accommodation in the compartment meeting the requirements for mass, centering
and inertial properties of the spacecraft and conditions for cargo delivery
(fastening, in-flight attitude, etc.), and the conditions stipulating that the equipment
in the compartments shall not interfere with the crew operating controls and using
life support equipment;
3)
physical integration – building up kits, accommodating and securing
cargoes in the spacecraft compartment.
6.6.3 Cargo accomodation
6.6.3.1 General
The cargos are arranged within the compartments in such a way as to meet the
requirements and constraints of the vehicle and of the cargoes themselves, and this
arrangement is covered in the vehicle design documentation.
Cargoes are installed onboard the spacecraft using RSC Energia drawings and
design and operational documentation contingent upon availability of certificates
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for transportation and operation, and the installation is covered in the onboard
documentation for the vehicle.
All the operations to service and process the deliverable cargoes at the processing
facility are only conducted subject to availability of proper instructions agreed with
RSC Energia and under the supervision of its responsible representatives.
6.6.3.2 Conditions for placing cargoes onboard Progress spacecraft
The owner/curator of the cargo shall keep in mind that after the cargo is stowed
in the vehicle and until it is unstowed onboard the station, there will be no access
to it. The duration of the above period depends on the conditions for stowing and
installing cargoes onboard the vehicle.
Deliverable cargoes are stowed in containers designed for them and in free areas
within the cargo compartment.
As a rule, cargoes with a mass of 8÷10 kg are stowed in the spacecraft containers
(if there are no special transportation requirements), and cargoes with a higher
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mass are installed in special transportation racks, for which purpose the deliverable
unit shall have mounting fixtures.
The need to develop special mounting fixtures and packaging (adapter racks, shock
absorbers, thermal insulation, moisture-proofing, etc.) is stipulated when the
contract is concluded.
6.6.3.3 Cargo accommodation onboard Soyuz spacecraft
Deliverable cargoes are stowed in the areas within the cargo compartment that
are set aside for them.
As a rule, cargoes with a mass of up to 5 kg are stowed in the spacecraft
containers (if there are no special transportation requirements), and cargoes with a
higher mass are installed in special transportation racks, for which purpose the
deliverable unit shall have mounting fixtures.
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The need to develop special mounting fixtures and packaging (adapter racks, shock
absorbers, thermal insulation, moisture-proofing, etc.) is stipulated when the
contract is concluded.
The main place for accommodating payloads returning to Earth is the container
under the middle seat. In the two-seater version of the spacecraft, returning
payloads are also stowed in a special container installed into the right-hand seat.
Returning payloads shall be capable of being loaded through the payload
container cover (dimensions 170x470 mm). The decision on the feasibility of
returning a cargo to Earth is made based on the results of RSC Energia study.
6.6.4 Deliverable cargo integration timeline
6.6.4.1 Deliverable cargo integration into Progress spacecraft
The work to integrate cargoes into Progress spacecraft starts 6 months before the
scheduled launch of the spacecraft and is based on the cargo delivery Manifest.
6 to 4 months before the spacecraft is launched, cargo outline and installation
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drawings are submitted to RSC Energia for approval.
Based on the results of the approval process for the cargo outline drawings, the
feasibility of delivering cargo onboard the spacecraft is determined, and cargo
delivery Manifest is updated (if necessary).
4 months before the spacecraft is launched, final outline and installation
drawings of the deliverable cargos are to be submitted to the RSC Energia
specialists on the spacecraft.
3 months before the launch, the specialists on the vehicle publish layout
drawings for cargo accommodation inside the cargo compartment of the spacecraft,
and perform a preliminary engineering analysis of mass, center of gravity, and
inertial characteristics of the spacecraft.
Published 1.5 months before the launch of the spacecraft are engineering
drawings and documentation on cargo installation, and the results of the analysis of
cargo mass, inertial, and center of gravity properties of the spacecraft.
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No later than 1 month before the launch, the cargo is to be delivered to the
processing facility, where it will then be physically integrated into the vehicle.
Fig. 6.6.1 shows the process of deliverable cargo integration into Progress
spacecraft.
L-4
L-6
L-3
L-1.5
L-1
L
Months
Cargo data are submitted and approved
Layout drawing is developed for
cargo accommodation in the spacecraft
Engineering documentation
and drawings are developed
Cargo delivery to the engineering
À
Physical integration of the cargo
À
Spacecraft launch
Fig. 6.6.1 The process of deliverable cargo integration into Progress
spacecraft.
Physical integration of deliverable cargos.
The process of physical integration of cargoes is determined by the technical plan
of spacecraft processing at the processing facility and the launch pad. Fig. 6.5.2
shows the process of physical integration of cargos into Progress spacecraft.
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facility
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L-25
L-12
L-8
L-5
L-2
L-1
L
Days
Loading the main load of deliverable cargoes
Loading additional deliverable
cargos
Loading additional cargo at the launch pad
À
The spacecraft launch
Fig. 6.6.2 The process of physical integration of cargo into Progress
spacecraft.
The major portion of physical integration of deliverable cargo into the cargo
compartment of the spacecraft is performed 25÷12 days before the launch of the
spacecraft. Specific schedules for loading cargo within the specified interval of
time and the number of days allocated for physical integration is determined by the
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technical plan for specific spacecraft processing.
Within the time period of 8÷5 days before the launch of the spacecraft, during
final operations after the spacecraft has gone through the filling station and
payload unit assembly, it is possible, by the decision of the technical management,
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to perform additional physical integration (additional loading) of deliverable
cargoes of a limited mass and volume.
2÷1 days before the spacecraft launch, it is possible, by the decision of the RSC
Energia General Designer, to load additional kits of a limited size through the side
loading hatch on the spacecraft. The need to integrate cargo during that time period
shall be stated when the cargo delivery request is submitted for inclusion in the
Manifest.
6.6.4.2 Deliverable cargo integration into Soyuz spacecraft
The work to integrate cargoes into Soyuz spacecraft starts 4 months before the
scheduled launch of the spacecraft and is based on the cargo delivery Manifest.
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4 to 3.5 months before the spacecraft is launched, cargo outline and installation
drawings are submitted to RSC Energia for approval.
Based on the results of the approval process for the cargo outline drawings, the
feasibility of delivering cargo onboard the spacecraft is determined, and cargo
delivery Manifest is updated (if necessary).
3 month before the spacecraft is launched, final outline and installation drawings
of the deliverable cargos are to be submitted to the RSC Energia specialists on the
spacecraft.
2.5 to 2 months before the launch, the specialists on the spacecraft publish layout
drawings for cargo accommodation inside the cargo compartment of the spacecraft,
and perform a preliminary engineering analysis of mass, center of gravity, and
inertial characteristics of the spacecraft.
Published 1.5 months before the launch of the spacecraft by specialists on the
spacecraft are engineering drawings and documentation on cargo installation, and
the results of the analysis of cargo mass, inertial, and center of gravity properties of
No later than 1 month before the launch, the cargo is to be delivered to the
processing facility, where it will then be physically integrated into the vehicle.
Fig. 6.5.3 shows the process of deliverable cargo integration into Soyuz
spacecraft.
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the spacecraft.
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L-4
L-3
L-2
L-1.5
L-1
L
Months
Cargo data are submitted and approved
Layout drawing is developed for
cargo accommodation in the spacecraft
Drawings and technical
Documentation is developed
À
Cargo delivery to the processing
facility
Physical integration of the cargo
À
Spacecraft launch
Physical integration of cargoes into Soyuz spacecraft
Fig. 6.5.4 shows the process of physical cargo integration into Soyuz spacecraft.
L-12
L-30
L-5
L-2
L-1
L
Days
Loading additional deliverable
cargos
Loading additional cargo at the launch pad
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L-8
Loading the main load of deliverable cargoes
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Fig. 6.6.3 The process of deliverable cargo integration into Soyuz spacecraft.
À
The spacecraft launch
Fig. 6.6.4 The process of cargo integration into Soyuz spacecraft
The major portion of physical integration of deliverable cargo into the orbital
module and the descent vehicle of the spacecraft is performed 30÷12 days before
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the launch of the spacecraft. Specific schedules for loading cargo within the
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specified interval of time and the number of days allocated for physical integration
is determined by the technical plan for specific spacecraft processing.
Within the time period of 8÷5 days before the launch of the spacecraft, during
final operations after the spacecraft has gone through the filling station and
payload unit assembly, it is possible, by the decision of the technical management,
to perform additional physical integration (additional loading) of deliverable
cargoes of a limited mass and volume.
2÷1 days before the spacecraft launch, it is possible, by the decision of the RSC
Energia General Designer, to load additional kits of a limited size through the side
loading hatch on the spacecraft. The need to integrate cargo during that time period
shall be stated when the cargo delivery request is submitted for inclusion in the
Manifest.
6.6.4.3 Safety documentation submittal deadlines
Deadlines for submitting transportation safety data packages for review are:
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−3 months before launch – for the first category cargoes (non-hazardous cargoes)
and second category cargoes (the cargoes that are hazardous, but prevention
measures are obvious);
−4 months before the launch – for the first and second category cargoes, which
require developing special fixtures for securing them onboard the spacecraft.
- 6 months before the launch – for the cargoes of the third category (highly
technical cargoes posing a great hazard during delivery, installation and operation),
a transportation safety data package containing all the currently available
information on potential hazards of the cargo must be submitted. The final
complete transportation safety data package shall be submitted within the
timeframe established for the first and second category cargoes.
Note: Deadlines for submitting transportation safety data package for review are
different from deadlines for submitting inputs for assessments and analyses.
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6.6.5 Integration timeline for cargoes to be disposed of
Considering the need to analyze and approve documentation packages for
cargoes delivered to ISS onboard the transportation vehicles of the Partners, cargo
documentation packages shall be submitted to RSC Energia from the cargo owners
(curators):
− for approval – 1 month before spacecraft undocking from the space station;
− approved – 15 days before the undocking.
No documentation packages need to be submitted for the cargoes that were
delivered to ISS onboard Russian vehicles.
6.6.6 Integration timeline for cargoes to be returned to Earth
Defined at 3 months before spacecraft undocks from the space station is: which
cargoes to be returned to Earth are to be integrated into the descent vehicle of the
spacecraft in orbit.
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Defined at 2.5 to 2 months before spacecraft undocks from the space station is
the configuration of the hardware to be returned in the order of priority. Submitted
to RSC Energia shall be final documentation packages only for those cargoes,
which were delivered to ISS onboard the transportation vehicles of the Partners.
1.5 to 1 month before the spacecraft undocking, RSC Energia specialists on the
vehicle define the layout for cargoes to be returned to Earth. These materials are
the basis for updating onboard documentation. The configuration and
accommodation of the equipment to be returned can be updated using a radiogram
from MCC after the Soyuz spacecraft has docked with the space station, but no
later than 5 days before the scheduled landing of the descent vehicle (provided the
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cargo safety package is available) in consultation with RSC Energia.
Fig. 6.6.5 shows the process of integrating cargoes to be returned to Earth into
Soyuz spacecraft.
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T-
T-
T-
T-
T-
75
60
45
30
21
T-7
T
Days
Submission of the manifest and the cargo data package
Preliminary design update
(layout configuration)
Updating onboard
documentation
Uplinking a preliminary radiogram from
MCC
preparing cargoes for return
Physical integration of the cargo to be returned into
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the descent vehicle
Docking of the next spacecraft
À
À
Descent vehicle landing
Fig. 6.5.5 The process of integration of cargo to be returned to Earth into
Soyuz spacecraft
6.6.7 Documentation requirements
6.6.7.1 Requirements for the outline and installation drawing
The outline and installation drawing is a drawing of a cargo (a kit, a cable) in
transportation position.
The outline and installation drawing shall cover the following data:
1)
name;
2)
drawing number;
3)
mass;
4)
coordinates of the center of mass (with tolerances);
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5)
moments of inertia (with tolerances);
6)
unit drawing/configuration;
7)
outline dimensions (with tolerances) of the cargo, specifying
dimensions of all
protruding elements;
8)
locations of connectors, monitoring or control panels, toggle switches;
9)
coordinates of the unit attachment points and dimensions of fastener
holes (for cargoes with a mass of more than 8÷10 kg.);
10)
instructions for the unit orientation during transportation onboard
spacecraft with respect to the flight/g-load direction, if there is such a
constraint (subsection 0);
0H
11)
instructions concerning the feasibility of the unit delivery inside a
spacecraft container or on-the-spot accommodation;
12)
special requirements for transportation conditions onboard the
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spacecraft (if need be);
13)
lifting points (for cargoes with a mass above 20 kg).
14)
bar-code location
If, during transportation onboard the spacecraft, the cargo is staying inside its
packaging or it has temporary protective elements (covers), the outline and
installation drawing shall reflect the cargo in packaging and with its protective
covers.
Outline and installation drawings of the cargoes shall be approved by specialists
on the spacecraft (RSC Energia).
In case of repeated/multiple deliveries of a cargo that has already been delivered
on the same type of spacecraft in the past, no additional approval of the outline and
installation drawing is required.
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6.6.7.2 Requirements for the safety data package
In accordance with SSP 50146, cargo safety information shall be presented in the
form of a safety data package. General requirements for the safety data package
contents and format, as well as cargo classification by category are defined in SSP
50146, Attachment D.
In order to certify cargoes for compliance with the transportation environments
and transportation safety requirements, the owners of these cargoes need to prepare
a transportation safety data package to reflect compliance with requirements of
Section 2.5.
Information about cargo hazards is considered complete, if each hazard in the
transportation safety data package has:
- hazard description;
- measures to prevent the hazard;
- methods of safety verification and documentation with verification results.
Cargo ground processing safety is reviewed separately from transportation
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safety, therefore it shall not be covered in the transportation safety data package.
6.6.7.3 Requirements for accompanying documentation
Accompanying documentation is submitted to RSC Energia in Russian and in
English.
Accompanying documentation contains:
- technical data sheet;
- cargo processing instructions (if need be);
- instructions for the crew (if need be).
The list of accompanying documentation for an active cargo is TBD (in
consultation between RSC Energia and the country supplying the cargo).
The technical data sheet contains the name, the drawing number, the serial
number, the useful life, as well as confirms that cargo has been fully processed and
is operational (verification report), complies with specifications and is ready for
flight operation within the ISS. The technical data sheet is published for each item
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on the deliverable cargo list, if several identical cargoes are delivered, one
technical data sheet specifying their quantity is published.
Fig. 6.6.6 shows the reference coordinate system of the spacecraft.
Fig. 6.6.7 shows the cargo compartment of the Progress spacecraft.
Fig. 6.6.8 shows the arrangement of containers in the cargo compartment of the
Progress spacecraft.
Fig. 6.6.9 shows dimensions of cargo containers on the Progress spacecraft.
Fig. 6.6.10 shows the container and locations for accommodating additional
cargo inside the orbital module of Soyuz-TMA spacecraft.
Fig. 6.6.11 shows cargo container in the descent vehicle of the Soyuz-TMA
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spacecraft.
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Fig. 6.6.6 Reference coordinate system of the spacecraft
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Center of mass
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Docking unit
Access hatch
Docking unit
cover envelope
Access hatch
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Loading hatch
Payload
containers
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Payload
containers
Fig. 6.6.7 Cargo compartment of the Progress spacecraft.
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Tier
Tier
Tier
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Container numbering scheme
Fig. 6.6.8 The arrangement of containers in the cargo compartment of the Progress
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spacecraft.
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Tier III containers
(No. III-1, III-2, III-3, III-4)
Tier II containers
(No. II-1, II-2, II-3)
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Tier I containers
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Spacecraft assembly
installation envelope
Spacecraft assembly
installation envelope
Fig. 6.6.9 Dimensions of cargo containers of the Progress spacecraft
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Payload
accommodation
area
a closet
a closet
Payload rack
a couch
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payload
accommodation area
a couch
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Additional
cargo rack
Note:
1. In containers No. 1,5,7,9 the cargo to be delivered is placed in addition to the standard
equipment.
2. The payload is only placed in container No.6, when a two-seater version of the of the
spacecraft is used.
Fig. 6.6.10 The container and locations for accommodating additional cargo inside
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the orbital module of the Soyuz-TMA spacecraft
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Container cover
(in open position)
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Container cover
(in open position)
Fig. 6.6.11 Cargo container in the descent vehicle of the Soyuz-TMA spacecraft.
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Attachment A
(obligatory)
List of Acronyms
АФУ
БИЛ
БИТС
БКС
Бортовая кабельная сеть
БНО
Баллистико-навигационное
обеспечение
БО
Бытовой отсек
БСР
Блок сопряжения
Подпись и дата
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Инв. № дубл.
Подпись и дата
БСР-ТМ
Инв.№ подл.
Антенно-фидерное
устройство
Бортовые информационные
листки
Бортовая информационнотелеметрическая система
Блок сопряжения
телеметрических массивов
БТА
Базовая точка активная
БТП
Базовая точка пассивная
БТС
ВЗП-У
ВнеКД
Базовая точка для полезной
нагрузки
Платформа виброзащитная
универсальная
Внекорабельная
деятельность
AFD
Antenna and Feeder Device
Onboard information sheets
OI TS
Onboard information and
telemetry system
Onboard cabling system
OM
Trajectory and navigation
support
Orbital Module
IU
Interface Unit
TM IU
Telemetry array interface
unit
Active Basepoint
Passive Basepoint
Payload Basepoint
EVA
Multipurpose vibration
isolation platform
Extravehiclular Activity
ВЧ
Высокая частота
High frequency
ГА
Гермоадаптер
Pressurized adapter
Галактические космические
лучи
Galactic cosmic rays
ГО
Головной обтекатель
Payload fairing
ГрО
Грузовой отсек
Cargo compartment
ИС
Информационная система
Information system
КЭ
Космические эксперименты
Space experiments
КЦН
Комплекс целевых нагрузок
ГКЛ
ЛИ
Летные испытания
MPF
Mission Payload Facility
Flight tests
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ЛПЭ
МИМ2
МИМ1
МЛМ
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Взам. инв. №
Малый исследовательский
модуль 2
Малый исследовательский
модуль 1
Многофункциональный
лабораторный модуль
MRM2 Mini Research Module 2
Микро метеоритная защита
МСС
Модуль сбора сообщений
Научная аппаратура
Linear energy transfer
MRM1 Mini Research Module 1
MLM
Multi-purpose Laboratory
Module
Micro-meteorid shield
Message collecting module
SE
Scientific Equipment
НИП
Наземный измерительный
пункт
Ground tracking station
НШС
Нештатная ситуация
Off-nominal situation
ОСК
Орбитальная система
координат
ПГО
Приборно грузовой отсек
OCS
ПН
Пакет данных по
безопасности
Программно-математическое SW
обеспечение
PL
Полезная нагрузка
ПО
Программное обеспечение
ПДБ
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LET
ММЗ
НА
ПМО
Orbital Coordinate System
Instrumentation Cargo
Compartment
Safety data package
Software
Payload
Software
Test program
ПЭИ
Программа
экспериментальных
испытаний
РО1
Рабочий отсек
Working compartment
Протоны радиационных
поясов Земли
Российской сегмент
ISS RS
РС МКС международной космической
станции
Радиотехническая система
РСУС
управления и связи
Радиоэлектронная
РЭА
аппаратура
РПЗ
СА
Инв.№ подл.
Линейная передача энергии
Спускаемый аппарат
Earth radiation belt protons
International Space Station
Russian Segment
RF command and
communications system
RF equipment
Descent vehicle
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СБИ
Система бортовых
измерений
СМ
Служебный модуль
SM
Onboard measurement
system
Service Module
СО1
Стыковочный отсек
DC1
Docking Compartment
СОТР
СТТС
СЭП
ТБУ-В
ТБУ-Н
ТВС
ТГК
Инв.№ подл.
Термостат
биотехнологический
универсальный
высокотемпературный
Термостат
биотехнологический
универсальный
низкотемпературный
УФП
ФГБ
ЦА
ЦИ
ЦУП-М
MBLT
T
Multipurpose
Biotechnological LowTemperature Thermostat
Тяжелые заряженные
частицы
Транспортный грузовой
корабль
Heavy charged particles
Ressuply spacecraft
Phone
Ультракороткие волны
УСТТС
MBHT Multipurpose
T
Biotechnological HighTemperature Thermostat
TV-system
УКВ
УРМ-Д
Telephone and telegraph
communications system
Power supply system
Телевизионная система
Телефон
УРМ
Thermal control system
Система энергопитания
ТЛФ
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ТЗЧ
Система обеспечения
теплового режима
Система телефоннотелеграфной связи
Универсальное рабочее
место
Универсальное рабочее
место доставляемое
Устройство сопряжения с
системой
Устройство фиксации
пассивное
Функциональный грузовой
блок
Ultra-short waves
MPW
Multi-Purpose Workstation
MPWD
Multi-Purpose Workstation
Deliverable
System interface device
Passive restraint
FGB
Mission payload
Целевая аппаратура
Payload data
Целевая информация
Центр управления полетами
Москвы
Functional cargo module
MCCM
Mission Control Center –
Moscow
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ШК
Airlock
Шлюзовая камера
ERA
European robotic arm
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Европейский манипулятор
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modified
replaced
new
deleted
Total
number of
pages in the
document
Document
number
Incoming
number of
accompanyin
g
documentati
on and date
Data
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Revision
Page numbers
Signature
REVISION LOG
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